Pratt & Whitney The Aircraft Engine and Its Operation_Rev1949_BZ

THE AIRCRAFT ENGINE AND ITS OPERATION

INSTALLATION ENGINEERING JANUARY 1949

First Printing, April 1946 Reprinted Revisions, May 1947 Reprinted u Revisions, January 1949

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TABLE OF CONTENTS

THE AIRCRAFT ENGINE AND ITS OPERATION

TABLE OF CONTENTS FOREWORD

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PART I THE AIRCRAFT POWERPLANT THE AIRCRAFT POWERPLANT .........................................,................. THE POWER SECTION........................................ . . . . . ............. INSTRUMENTS AND CONTROLS...................................................... THE ACCESSORY REAR SECTION, THE NOSE SECTION, AND THE PROPELLER .................................... THE LUBRICATION SYSTEM ............................. .. .. . .... .. . THE COOLING SYSTEM ........................... ..................... ..... THE INDUCTION SYSTEM..................................... ................................. 1. Carburetion. .................................................... . ..... . 2. Supercharging ...;.................................................. ,,, 3. Ducting ...................................................... ................ ., . . THE IGNITION SYSTEM..................................................... .

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11 15 19 25

25 33 42

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PART I1 POWER, BMEP AND RATINGS POWER AND ITS MEASUREMENT .............................. BMEP ........... .....;..... ........................... ....................... . . RATINGS ............................... ......., ......... . ..

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PART, I11 GENERAL OPERATING INSTRUCTIONS

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................. ........ STARTING ............... ........................... .................+.. Control Position Check.............................................................. Clearing The Engine.................................. ..,.,,,....,..,.,........... Priming ..............., .............,........ ..,.......... ..... ........... Starting ................................................., ,.,,...,........,.........

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PRATT & WHITNEY AIRCRAFT

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GROUND OPERATION ...................................................................................... 84 Warm-up ................................................ ...................................... 84 Magneto Safety Check ..................................................................... 86 Ground Check .................................................................................... 87 Idle Mixture Setting ....................................................................... 90 Taxiing ............................................................:.................. ............ 91 COLD WEATHER OPERATION .................................................................. 92 Preheating ............................................................................. .............. 92 Starting Cold Engines .................:.................................................94 Warm-up ............................................................................................... 94 Flight Operation .............................................................................. 95 Dilution ................................................................................................... 95 FLIGHT OPERATION ..................................................................... .............. 97 Take-Off .................................................................................................97 Transition From Take-Off To Climb ....................................... 101 Flight - General ................................................................................ 102 Power . Selection ........................................................................102 Bmep Selection .........................................................................104 Altitude Selection ...................................................................... 108 Induction System Icing .........................................;.................... 109 Climb ..................................................................................................... 111 Transition From Climb To Level Flight .................................. 114 Level Flight ......................................................................................... 115 Descent ..................................................................................... .... 119 Glide and Approach .......................... . . ...................................120 122 Landing ................................... :......................................................... Stopping .............................................................................................. 122 Shutdown ;........................................................................................... 122

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PART IV ADVANCED SUPERCHARGING

FOREWORD ...................................................................................................... 124 PURPOSE OF SUPERCHARGING.............................................................124 134 RESULTS OF SUPERCHARGING............................................................ THE SINGLE-STAGE SUPERCHARGER .................................................. 135 1 Ground Boosting ...................................................................... 135 2. Low Altitude .................................................. ...................... 136 3. High Altitude ........................................................................... 137 Single-Stage, Two-Speed .............................................................. 138 Single-Stage, Variable-Speed ........................................................140 THE TWO-STAGE SUPERCHARGER ............. . . .................................... 141 1 Two.Stage. Two-Speed ...........................................................144 2 Two.Stage. Variable-Speed ................................................... 145 3 Other Two-Stage Systems .......................................'................ 146

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PRATI' & WHITNEY AIRCRAFT

FOREWORD

FOREWORD The aircraft engine is designed to power aircraft in flight. Its construction, installation, servicing, and repair present many interesting and important problems ; but its actual operation is, naturally, of more immediate concern to the pilot. A number of texts, handbooks, and manuals are available which discuss the principles of engine design and installation, and describe the procedures of maintenance and overhaul. As a supplement to this material, i t is the purpose of this book to discuss the powerplant as 'a functioning unit and an integral part of the aircraft in which it is installed, and so to provide a basis for an understanding of sound operating practice. The first part of the book is devoted to a description of the basic engine sections and "systems" which together make up the powerplant, and discusses them in relation to the different instruments and controls by means of which the pilot regulates the performance of his engine. The second part deals with the fundamental concepts of power and its measurement, bmep, and ratings, with particular reference to the principles of engine operation. The last part of the book consists of a set of general operating instructions in which the material treated in the first two parts is applied to the specific problems of engine operation encountered in flight and on the ground. I t is assumed that the reader is generally familiar with the main features of conventional engine design, with the names and functions of the more important units of the powerplant, and with the principles of the four-stroke cycle. In discussing the various temperature conditions of engine operation, a problem presents itself as to the relation between centigrade and Fahrenheit values. The Military Services and a considerable number of other operators use the centigrade system entirely, but a t the same time, a large group continues to rely on instrumentation that is graduated in fahrenheit. Because of its more prevalent use, Fahrenheit is the basic scale used in all discussions of this book. However, when the equivalent centigrade temperature is included the reader will observe that i t is not the exact equivalent of the Fahrenheit value ; for example, in mentioning a minimum oil temperature for operation 100 F (40C) is used. By this means the practical values in each system are recommended rather than an odd number which would not be easily carried in the mind, such as 100 F (38C).An exception to this rule are those temperature limits which are dictated by specifications governing the licensing or acceptance of the engine where obviously there can be no discrepancy between equivalent Fahrenheit and centigrade values. Many of the numerical values quoted are for purposes of illustration and may not apply to all engines. For specific values for any particular engine consult the applicable Specific Operating Instructions for the engine in question. The material in the book and its supplements relates principally to the radial, air-cooled engines manufactured by Pratt & Whitney Aircraft, and is based on the experience of more than two decades in operating these engines.

THE AIRCRAFT POWERPLANT

OPERATION & DESIGN

PART I

THE AIRCRAFT POWERPLANT OPERATION

The aircraft engine and the automobile engine are fundamentally alike in principle and design. Both convert the chemical energy of a mixture of fuel and air into useful work by the same process of internal combustion. Both incorporate the same basic mechanisms. Both furnish the motive power for the vehicles in which they are installed.

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The operation of the aircraft engine, however, differs in many ways from that of the automobile engine. The performance required of the latter is relatively limited, and weight may be generously used to insure wide margins of safety. Even the uninformed driver can do little to abuse his engine, and to this extent the automobile powerplant has been "foolproofed." By contrast, a wide range of performance is demanded of the aircraft engine, while the weight of the powerplant must necessarily be kept to a minimum in the interests of over-all efficiency of the airplane. Accordingly, the operator must substitute his understanding and judgment for mere weight and solidity of structure, and, in place of "foolproofing" the engine, the operator must be "fool~roofed" instead. Engine ratings, or the restrictions placed on engine operation, are of little practical interest to the automobile driver, since in the course of ordinary driving he cannot easily exceed these limits. On the other hand, engine ratings must be one of the first concerns of the airplane pilot. It 2s possible to operate virtually any aircraft engine a t excessive powers, and in most instances there is nothing to restrain the operator from exceeding the limits set by the ratings except his understanding and judgment. If the pilot makes use of engine speeds and cylinder pressures that are above the rated limits, he is relying upon margins of safety which may be dangerously narrow. If he fails to observe cylinder head or oil temperature limits, or operates

without regard to oil and fuel pressures or to fuel-air ratio, he is creating a situation that can only lead to lowered reliability and possible engine failure. In operating an aircraft engine the pilot is confronted with the basic problem of combining engine speed and cylinder pressure in order to obtain power. The manner in which these two factors are combined, together with the maintenance of suitable conditions of cooling, lubrication, and mixture strength, determines the efficiency of the performance and the dependability and durability of the power plant. DESIGN

Familiarity with the engine is basic to the understanding of sound operating practice. From an operator's point-of view, a typical air-, craft power plant may be considered as consisting of: 1. A Power Section 2. An Accessory Rear Section and a Nose Section, together with the Propeller 3. A Lubricating System 4. A Cooling System 5. An Induction System comprising Carburetion and Supercharging Systems 6. An Ignition System

A set of cockpit instruments, supplemented by the observed behavior of powerplant and airplane, enable the operator to determine the manner in which the different parts of his engine are functioning. A corresponding series d controls permits him to regulate their functioning. Efficient operation calls for the intelligent coordination of the several functions of the various sections and systems of the powerplant to meet changing conditions as they are encountered, having in mind the type of performance desired, and with due regard for the safety and protection of the engine.


PWA. 01. 100 /-'

THE POWER SECTION PARTS AND FUNCTIONS As the name implies, the power section is the section of the engine in which the power is developed, and which is the first to receive the impact of the loads created. I t consists basically of the following assemblies :

stroke ; imprison it during the compression and power strokes ; and release the burnt gases during the exhaust stroke. The valve-operating

VALVE SPRINGS

1. The Cylinders The aircraft engine is a heat engine which derives its power from the burning of a combustible mixture of gasoline and air. The fuelair charge is ignited by a spark, and the heat generated by the burning of the charge expands the gases formed during its combustion.

VALVE TAPPET VALVE TAPPET ROLLER

CYLINDER HEAD OIL DRAIN-

EXHAUST POR INTAKE POR SPARK PLUG BUSHING

mechanism consists of a set of cams and a series of tappets, push rods, and .rocker arms actuated by the cams to open and close each valve a t the proper moment in the power cycle. 3. The Pistons

COMPRESSION RINGS

COOLING F/NS

OIL CONTROL RING

PISTON PIN

OIL SCRAPER RING

MASTER OR ARTICULATED ROO

The chemical energy of the fuel and air is thus transformed into the mechanical energy of the expanding gases. Together with the pistons, the cylinders form the chamber in which the power-producing combustion takes place, and it is in them that the greatest concentration of heat and pressure is found. I

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2. The Valves and the Valve-Operating Mechanism The valves control the, admission of the fuelair charge t o the cylinders during the intake

The pistons receive the force of the expanding gases and transfer i t to the articulated and master rods. Like the cylinders, the pistons are subject to the heat and pressure of combustion, a s well as to reciprocating forces. 4. The Master and Articulated Rod Assemblies The master rod and articulated rods transfer the reciprocating motion of the pistons to the crankshaft where it is transformed into rotary motion. A t the piston ends the rods are subject to reciprocating forces and to pressure loads resulting from combustion, while a t the crankshaft centrifugal loads are also imposed on the

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THE POWER SECTION


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assembly as a consequence of the high rotational speed of the engine. There is one master

' E&PISTON PIN

BFAR/NG

ART/CULATED OR LINK RODS

KNUCKLE PINS

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and articulated rod assembly and one crank throw for each row of cylinders. 5. The Crankshaft The crankshaft receives the power developed in the cylinders from the master and articulated rod assembly, and transfers i t through gear

COUN TCRWEIGHTS DyNAM/ C DAMPER PROPELLER SHAFT ,WITH SPLINES

trains to the propeller, the supercharger, the valve-operating mechanism, the oil and fuel pumps, the magnetos, the propeller governor, the generator and other accessories. It is subject to a combination of centrifugal, reciprocating, and pressure loads, together with torsional stresses.

POWER SECTION CONTROL While conditions in the power section are ultimately the chief concern of the operator, the control of this section is not direct, and only one instrument, the cylinder head temperature gage, is actually connected to it. Instead, control is exercised through other parts of the powerplant, and the effects of this regulation are measured indirectly by various instruments which indicate conditions obtained in the nose and rear sections and in the lubricating, cooling, induction, and ignition systems. For example, cylinder pressures are controlled by the throttle of the induction system and by the propeller speed governor on the nose section; and these pressures are measured in terms of the readings of the manifold pressure and carburetor air temperature gages of the induction system (or the torquemeter gage from the nose section) and the tachometer connected to the rear section. An understanding of proper engine operation accordingly requires : 1. A familiarity with the various cockpit instruments and controls -their names and their location. 2. An appreciation of the meaning of the reading of each instrument--the condition i t indicates a t the particular part of the powerplant to which i t is connected, and the significance of this condition a s i t affects the power section and other parts of the engine. 3. A knowledge of the function of each control - the particular part of the powerplant which i t controls, and the effect of this regulation on conditions in the power section and elsewhere through the engine.

6. The Crankcase

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The crankcase, generally in two or more parts, houses and supports the crankshaft, and to it are attached the cylinders and the nose and supercharger housing sections. Its bearings absorb the power-producing loads from the pistons and crankshaft.

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THE ACCESSORY REAR SE,CTION


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THE ACCESSORY REAR SECTION, THE NOSE SECTION, AND THE PROPELLER TACHOMETER LEVER

REDUCTION GEAR

PROPELLER

QIL SCAVENGE PUMP

THE ACCESSORY REAR SECTION Crankshaft power is transmitted by means of an accessory drive shaft through the supercharger housing to the accessory rear section, where a series of gear trains, in turn, transmit the power to drive the oil pressure and scavenge pumps, .the magnetos (if rear-mounted), and various engine-driven accessories, including the TACHOMETER DRIVE

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fuel pump, the vacuum and hydraulic pumps, the generator, the tachometers, and the gun synchronizer. The accessory rear section houses the oil pressure and scavenge pumps; provides mounting pads for the magnetos and the engine-driven accessories; and furnishes connections f o r the starter, the auxiliary power drives, and the oil lines.

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TACHOMETER DRIVE - R.H. STARTER DRIVE

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VACUUM OR HYDRAULIC PUMP DRIVE

Accessory locations are typical and do not apply to all engines.

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THE NOSE SECTION

2. Geared Drive

Crankshaft power is transmitted forward through the nose section to the propeller. The nose section houses the propeller shaft reduction gearing and one or more oil scavenge pumps. It provides a power take-off for the attachment of

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a propeller or helicopter rotor drive; and i t furnishes mounting pads and drives for the scavenge pumps, the nose-mounted magnetos and distributors, the propeller governor, and other accessories.

REDUCTION GEARING DRIVE RATIOS

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A

Small engines which operate a t relatively low rpm and which are equipped with propellers of small diameter do not normally require a reduction gear between the crankshaft and the propeller. In such cases the propeller shaft is

an extension of the crankshaft,, and the propeller is directly clriven a t engine speed.

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Larger engines require bigger propellers to absorb the greater power developed. The limits of propeller size and rpm are reached when the velocity of the blade tips exceeds that of sound, a t which point the thrust efficiency of the propeller begins to break down, while the energy of the engine is dissipated in making noise. By placing a speed reduction gearing between the crankshaft and the propeller shaft, both engine and propeller can be made to operate a t efficient speeds. The range of reduction gear ratios is from 4:3 to 3:l.

THE AIRCRAFT POWLRPLANT -

THE PROPELLER The propeller is essentially a device for converting shaft horsepower into thrust horsepower. It is quite possible, however, for a propeller to absorb the entire power output of an engine and deliver little or no useful work in return. If i t is to function efficiently the propeller must be: 1.Correctly designed for its particular instal* lation - size, shape, number and pitch of blades, and freedom from excessive vibration. - 2. Operated within the proper speed range for its design.

as manifold pressure, remain unchanged, the rpm will decrease. Conversely, a decrease in pitch will result in a higher rpm. Engine power , and speed, accordingly, no longer bear any fixed relation to one another. The constant speed propeller is a variable pitch propeller in which the regulation of the pitch is automatically controlled by a governor in such a way that the rpm is maintained unchanged a t a figure selected by the pilot, regardless of changes in power or .manifold pressure. Such an arrangement offers the greatest possible flexibility of operation, but i t places upon the operator the responsibility of choosing a well-

PROPELLER TYPES Propellers may be divided into general types: 1. Fixed Pitch

The pitch, or blade angle, of this type of propeller is fixed and cannot be varied in flight. (In some models the pitch is adjustable on the ground.) This installation is light and simple, and gives satisfactory performance within a limited range of power and speed. The power absorbed by this type of propeller varies as the cube of the rpm - i.e., bhp = K X rpm3. Accordingly, for any given power there is a corresponding rpm. Since power can be regulated by manifold pressure only with this type of installation, propeller speed is controlled by the throttle. 2. Two Position A hydraulic mechanism, operated by engine oil pressure, makes i t possible for the pilot to select in flight either a high pitch (low rpm) or a low pitch (high rpm) position for more efficient speed regulation. A desired power may thus be obtained with two different combinations of rpm and manifold pressure : a high rpm and low manifold pressure or the opposite combination in high pitch. In both positions engine power and, hence, speed are regulated by the throttle, as with a fixed pitch propeller.

3. Variable Pitch

A hydraulic, electrical, or mechanical mechan-

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ism permits the pilot to set the propeller blades while in flight to any desired angle over a considerable range. As the pitch of the propeller is increased the resistance offered t o i t by the air becomes greater, and if other factors, such

balanced combination of rpm and manifold pressure to obtain a desired power and, a t the same time, make the most efficient use of both engine and propeller. , An extension of the variable pitch principle makes it possible to set the propeller blades roughly parallel to the line of flight, and so stop the engine while in the air. This is known a s feathering, and is often found useful in emergencies which involve engine failure. With the blades in the feathered position, the drag of the propeller is reduced, while the possibility of further damage to aircraft and powerplant from a "windmilling" engine is prevented. It is also possible with some types of propellers to turn the blades past full low pitch into a reversed pitch position, thus reversing the thrust of the propeller - a feature which is helpful in maneuvering a seaplane on the water or braking an aircraft on the ground.

RPM CONTROL Engine speed, or rpm, is thus controlled in one of the following ways:

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PWA. 01. 100

1. By the throttle lever, in the case of fixed pitch propellers (or the high or low pitch position of a two-position propeller) ; or, 2. By the propeller governor control lever, in the case of variable pitch propellers.

Constant speed propeller installations commonly include a governor driven by a gear train from the crankshaft and mounted on the nose section. By directing engine oil under high pressure to the mechanism contained in the propeller

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GOVERNOR

GOVERNOR FLYWEIGHT

OR TO HYDlPAULlC -ELL SYSTEM

hub, or by controlling the eleczrical impulses sent to a hub-enclosed motor, the governor regulates the pitch of the propeller blades in such a way a s to maintain a constant rpm regardless of power, manifold pressure, or altitude. On other installations the propeller itself includes the governor a s an integral part of the hub, and the complete control system, including the oil supply, is independent of the engine.

OIL ,PEL(

COcKPlr CONTROL

:SSu'RE PUMP OIL SUPPLY

It should be noted that, if the governor is set to regulate a t a high or a t maximum rpm, and the engine is not developing sufficient power to turn the propeller a t this speed, the latter will be driven in full low pitch in the same manner as if i t were a fixed pitch propeller. In other words, propeller speed is a function of manifold pressure, and is controlled by the throttle until the governing rpm is attained. This characteristic is made use of in pre-flight checks and other ground operations a t low power. Proper operation of hydraulic propellers demands that the oil be a t the correct temperature and pressure. Oil that is too cold to flow freely will result in a sluggish response of the propeller to the governor. In the case of electric propellers, proper operation requires that the electric power reserve be capable of delivering a current suffi: cient to actuate the pitch control mechanism. Operation with weak batteries drained by other demands can lead to serious over-speeding.

THE LUBRICATING SYSTEM


THE LUBRICATING SYSTEM Oil is often referred to a s the "lifeblood of the engine," and by analogy the lubricating system may be somewhat crudely compared with the human circulatory system. Oil is forced from the pressure section of the oil pump (right side of heart) through various pipes and passages (arteries) to different parts of the engine (capillaries). After performing its lubricating and cooling functions, the oil is collected in one or more sumps where i t is drawn through other passages and pipes (veins) by one of the several scavenge pumps (left side of heart), and sent thence through the oil cooler (lungs-the analogy is a trifle strained) to the tank, and so back once more to the pressure pump. The principal features of a typical lubricating system are shown, and their functions briefly explained, in Fig. 1 on page 16. FUNCTION

The lubricating system of an aircraft engine performs three functions :

If two such surfaces were to slide over one another in direct contact, the result would be an

iabrfemaee between the surface peaks, which weald l a d to excessive friction and end by destroying the surf-.

By separating the rurfaeea with a film of d l the friction is reduced to that existing between the mo~wuleeof the lubricant. The oil in d i m t

1. I t reduces friction between sliding and rolling surfaces.

To the naked eye a highly finished metal surface appears like this:

But under a m i a m o p e a cross-section wouM look more like this :

contact with the surfaces moves with the Burfaces; friction then occurs only by reason of the iatermediatx! oil layers sliding over one another. With perfect lubrication no wear of the bearing surfaces should weur erzept, po5~$bly,at starting, for at all other tjm& they are separated by an oil film. When properly designed and lubricated, plain bearings may have a coefficient of friction as low m d s load carrying capacity as high alir ball or m l k bearings.

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SPOlSH EZD TO VIBRATION DAMPENER

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LUBRICATING SYSTEM

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OIL

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OIL PRESSURE P

O/L SCAVANGE

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PRESSURE RELIEF

OIL PRESSURE R

SUMP SCREEN

MA/N OIL SUMP

OIL COLLECTED F

MTERED TRANSFE

THE LUBRICATING SYSTEM

THE AIRCRAFT POWERPLANT 2. The lubricating system cools parts of the engine that cannot be cooled by air.

CONTROL OF OIL PRESSURE, TEMPERATURE, A N D FLOW

For example: the piston, which forms one end of the combustion chamber, is subject t o the most intense heat present in the engine. To cool it a spray of oil is directed to the under side of the piston through jets located in the crankshaft. This spray also lubricates the cylinder walls and piston pins, and assists in cooling the

To perform its several functions satisfactorily, lubricating oil must be : 1. Within the proper pressure limits. If oil pressure is too low, oil may not reach all parts of the engine, will not be forced in between bear/5,000

LB

OR MORE LOAD

MASTER ROD MASTER ROD CRANKSMF T

lower parts of the cylinder wall. The master roa bearing is another example of a part that is both cooled and lubricated by the oil of the system. Approximately 10% of the heat released by combustion is removed from the engine through the lubricating system. 3. The lubricating system furnishes oil for the

operation of various units and accessories, such a s : a. The blade angle varying mechanism of the hydraulically-operated propeller. (In this instance engine oil pressure is boosted by the propeller governor or by the feathering pump.) b. The impeller drive clutches, variable spark advance mechanism, manifold pressure regulator, and other units. (Units such as these are operated by normal engine oil pressure.)

If oil pressure is too high, oiI leaks are likely to develop and oil consumption to become exeessiw . Of the various cockpit gage8, none is more important than the oil pressure gage. Loss of oil pressure is an indication of the failure of the oil supply, and without lubrication i t may he only a matter of seconds before parts have heated and burned, bearings have seized, and the engine is brought to an abrupt halt, badly damaged.

This emphasizes the vital importance of making the proper oil pressure checks prior to takeoff. An oil pressure relief valve fixes an umer

In performing its lubricating and cooling functions the oil absorbs a considerable amount of heat-the eauivalent of 30 horsepower or more

cold oil, pressures of several hundred psi may be encountered when starting.) The relief valve must be set on the ground, and cannot be adjusted in flight.

cases and the surfaces of the oil tank; on larger engines an external oil cooler is necessary. The oiI cooler is equipped with shutters or flaps by means of which the flow of cooling air can be regulated to obtain the desired oil temperatures. The oil cooler shutters are controlled automatically or manually by means of a cockpit lever.

2. Within the proper temperature limits. If the

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I t is the practice to measure the oil temperature a t a point in the engine lubrication system that is close to the oil inlet. By taking the temperature a t this point, the oil temperature reading is then a check of the condition of the lubricant before i t enters the engine. Measuring the oil temperature a t the outlet has been seriously considered but, a s the temperature a t this point is a measure of the oil temperature rise in the engine, the gage reading will vary widely with changing conditions and will not constitute a stable standard for determining the condition of the oil. If, during a climb with wide open shutters, the airflow does not provide sufficient cooling, a greater flow of air may be obtained by reing oil temperature is to reduce power. This should be done by lowering the rpm as well a s the manifold pressure, since friction and oil temperature are affected by engine speed even more than by power.

3. Of suitable quality and sufficient mechani cal strength so that the oil film will not breal down under any condition of temperature an( pressure likely to be encountered.

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CONTROL OF THE LUBRICATING SYSTEil

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IIf the oil is too hot. i t cannot s u ~ ~ oheavy rt

The operator has no control of the lubricatin system while in flight other than to regulate t h temperature of the oil in response to the real

consequence, oil pressure may drop below acceptable limits, and oil consumption becomes excessive.

ders i t imperative that all ground checks b scrupulously performed to make certain that a adjustments have been properly made.

oil is too cold, i t will not flow freely.

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PRATT & WHITNEY AIRCRAFT (232-260 C ) * , the material will be seriously weakened.

2. Damage t o cylinder and piston parts. Excessive temperatures may lead t o warping of the valves and valve seats, and failure of valve stem and rocker arm lubrication, and the breakdown of the oil film between piston and cylinder, with the possibility of scoring and even seizure. 3. Increased tendency t o detonation. High temperatures also increase the tendency of the fuel-air charge to detonate. (See pp 27-29) Detonation in turn causes high cylinder head temperatures - and so a vicious cycle is begun. Regardless of which comes first, however, the effect on the engine is inevitably damaging.

PWA. 01. 100 The thermocouple location is selected because of i t s stability as a temperature reference and because of the mechanical advantages for making a practical and durable connection. The temperature recorded a t either of these points is merely a reference or control temperature; but so long as it is kept within the prescribed limits, numerous tests have shown t h a t the temperatures of the cylinder dome, the exhaust valve, with its seat and guide, and the piston will also be within a satisfactory range. As a rough example, it can be imagined t h a t the household thermostat has been placed above

TEMPERATURE INDICATION Cylinder head temperatures are indicated by means of a gage connected to a thermocouple attached t o the cylinder which experience shows t o be the hnt,test on Ecn engine in any particular installation. The thermocouple itself is located :

Under a rear spark plug in a special gasket. a radiator. In order to obtain the desired room temperature it will be necessary to set the thermostat (reference or control temperature) to some higher figure. The thermocouple is t h e cylinder's thermostat indicator. Its limit is established, arbitrarily, in order to keep a t satisfactory temperatures the vital portions of the cylinder whose temperature cannot be measured directly. In a special well in t h e top or rear of the cylinder head.

* The cylinder temperatures quoted above and in the remaining portions of this book are those which have applied to all engine models heretofore manufactured by Pratt & Whitneg Aircraft. Limits applying to engines of other manufacture, or to future Pratt & Whitney Aircraft models, may differ somewhat from the temperatures quoted above.

As the thermocouple is attached tr, one cylinder only i t can do no more than give evidence of general engine conditions. While normally i t can be presumed t h a t the remaining cylinder temperatures will be lower, local conditions such a s detonation or ignition failure will not be indicated unless they occur on the therrnocoupIe cylinder.

THE CYLINDER COOLING SYSTEM

THE AIRCRAFT POWERPLANT LIMITING TEMPERATURE Two limiting cylinder head temperatures are usually found in specific engine operating instructions : 1. The higher limiting temperature is for a restricted period of time, and is confined to take-off, t o maximum performance in climb and level flight, and to emergencies.

head temperature 50 F (30 C) below this limit to keep the cylinder head materials a t high operating strength. THE COOLING SYSTEM

e; I

In all aircraft cooling systems air is use carry the heat away from the cylinders. When this is done directly, without an intermediate liquid coolant, the problem of cooling largely resolves itself into that of:

1. Exposing a sufficient surface area of the cylinders to the cooling airflow. 2.

ABSOLUTE

MAXfMUM

FOR

@#€*%/OUR

ONLY

The temperature limit for restricted operations should, therefore, be used for the shortest possible time only, and must never be exceeded. 2. The lower limiting temperature is the maximum for continuous operation. I t should never

BES J OPERAT/ON

I

MAXfMUM FOR JINUOUS OPERATION

be exceeded except under the restricted operating conditions mentioned in the previous paragraph. I t is sound practice to hold t h e cylinder

Directing the air efficiently against all parts of the cylinders.

3. Providing a sufficient cooIing airflow, together with some means of regulating the airflow in response t o varying conditions.

The surface area of the cylinders exposed to the cooling airflow is increased by the use of cooling fins. The efficiency of t h e fins is increased, up to a certain point, a s they are made deeper and spaced more closely together. Not many years ago engine cooling was accomplished by simply holding the cylinders into the breeze. This arrangement had several shortcomings. While all the air in the path of the engine passed between the cylinders of the engine, only a small portion of i t came close enough to the fins to remove much of t h e heat. The cylinders were cooled unevenly, the rear portions of each one tending to run hotter than the front. The turbulence created affected all parts of the airplane a f t of the engine.

'RATT & WHITNEY AIRCRAFT

--3Fh

PWA. 01. 100

+--DtX.EcToR BAFFLE

More recently inter-cylinder and cylinder head baffles have been used to force the cooling air into close contact with all parts of the cylinders, In this way the entire airflow is used to carry the heat away from the fins. Although the resistance offered by the baffles and fins to the passages of the cooling air demands t h a t an appreciable pressure differential be maintained across the engine to obtain the necessary air-

-n'F="

BAFFLE SEAL

flow, the volume of cooling air required is much reduced by the use of a properly designed system of deflectors, and the resulting drag is actually decreased. A flow of cooling air sufficient to carry away the heat from the cylinders is obtained by means of the cowl which encloses the engine. The flow of cooling air may be regulated by a series of adjustable cowl flaps a t the rear of the cowling.

The pressure differential between the front ( 2 ) , and rear (3), of the engine forces the air past the cyiinders through the passages formed by the baffles, deflectors, and fins.

As it nears the cowl exits (4), the air is speeded up by restricted passages to merge smoothly with the

air stream.


Actually, only about I U ~01 tne air in the path of the engine is forced through the cowl, the remainder passing smoothly over the outside of the cowling. As a result there is little cooling drag-the power consumed in cooling amounts to roughly 57; of the bhp-and performance is not sacrificed because of the shape of the engine and its cowling.

TEMPERATURE CONTROL The amount of heat transferred to the cooling air is roughly proportional to the mass airflow. Control of engine temperatures may, therefore, be accomplished by regulating the number of pounds of air that are forced past the engine in a given time by the cowling. Mass airflow in turn is regulated by varying the pressure drop across the engine. This may be done in one of two ways : 1. By Cowl Flaps

Fixed exit cowls are generally satisfactory on aircraft whose range of speed is small, so long as the exit gap is sufficiently large to provide the necessary airflow with the entrance pressure available a t the lowest climbing speed. At the higher speeds of level flight the increase

THE CYLINDER COOLING SYSTEM

in pressure a t the cowl entrance forces more air past the engine than is actually required; but, as the speed range is low, the small waste of energy involved does not warrant the extra weight and complication of adjustable cowl flaps. On the other hand, with high performance aircraft there is an appreciable waste of power in forcing excess cooling air past the engine a t high speeds, and the extra weight and complication of adjustable cowl flaps are accordingly justified. At high speeds the exit is reduced so that only the required airflow can pass through. At low speeds the exit area is increased. This decreases the pressure behind the cylinders to compensate for the reduced frontal pressure, and so maintains the necessary drop across the engine. Cowl flaps thus act a s a throttle for the cooling airflow. The effective range of cowl flap opening is limited by aerodynamic considerations, which vary from installation to installation. For example: the parasite, or form drag of the aircraft is increased as the flaps are opened; and, if the flaps are opened too far, buffeting may result above certain airspeeds, and the control of the aircraft consequently disturbed.

.

I

PRA'IT & WHITNEY AIRCRAFT

2. By Air Speed Any variation in the forward speed of the airplane affects the p r m u m a t the face wf the engine and, hence, the mass airflow past the cylinders. If airspeed can be increased without calling upon the engine for additional powerfor example, by decreasing the rate of climb-the result will be a more effective cooling of the engine.

PWA. 01. 100

, k b w this fwl-%~!ir ratio the exeess of air in the charge r e d w a , the earnbustion temperature. When the mixture Is richer than this fuel-air ratir~the presence of excess fud reduces the

There are also two ways of controlling the temperature of an engine, neither af which, strictly speaking, is effected by means of the powerplant's cooling system.

eombustian temperature. A measure of ccmtrol over cylinder head temperatures can then be obtained by inweasiag the fuel-air ratio from sutmi3c lean ta automatic rich p m v i M that the automatic rich mixture is sufficiently richer than best power. Generally, in the cruising m g e * advancing the mixture control from automatic lean to au&etticl rich will raise the mixture strength Pmm near best economy up to agproximately lean best power with the remalt that the e ~ l i n d whead temperature will be imremed rather than lowered.

1. Fuel-Air Ratio.

2. Psweir.

The heat generated by combustion is g r e a t e ~ t with a fuel-air ratio of approximately 0.067 slightly below the so-called Iean best power mixture setting. (See page 26). If the mixture is

A redudion in power will also ewse a drop in cylinder h a d temperatures. To obtain the greatest effwt, power should be reduced by

OTHER TEMPERATURE CONTROLS

lowering Bokh rp& and manifold pressure.

THE INDUCTION SYSTEM-1.

THE AIRCRAFT POWERPLANT,

CARBURETION

INDUCTION SYSTEM FUNCTION The function of the induction system is to deliver a combustible mixture of fuel and air to the cylinders and, subject to intelligent control by the operator, to deliver the charge: (a) a t the proper fuel-air ratio for the particular type of operation demanded ;

The induction system can conveniently be discussed under three general heads: 1. Carburetion

- dealing principally

with the carburetor and the means of obtaining the desired fuel-air ratios.

(b) in sufficient quantities to meet all power requirements ;

2. Supercharging - which is chiefly concerned with providing an adequate airflow without too great a temperature rise or power restriction.

(c) a t a temperature that will insure efficient combustion ;

3. Ducting - which deals with the engine's air conditioning system.

1. CARBURETION THE FUEL-AIR RATIO The determination of the combustible range of a gasoline-air mixture is a simple and sporting investigation: The first step is to light a match in air containing no gasoline vapor. The match will burn, but there will be no explosion-proving that air merely supports combustion. The second step is to remove carefully all the fuel vapors above an open dish of fuel, and then dash a lighted match into the liquid. If the experimenter succeeds in getting the lighted match into the gasoline, the flame will be extinguished -proving that gasoline by itself is not combustible. More than likely the investigator will not

I

MIXTURF 7-00 LEAN TO PRODUCE POWER 7

have completely removed the gasoline vapors, and he will have succeeded in getting the burning match only part way to the fuel when he will be mad-e painfully aware of the forces that act on the piston of an internal combustion engine -proving that a mixture of gasoline vapor and air can be violently explosive, and that home experiments should be rigidly controlled. The variations in energy with different mixtures of gasoline and air are illustrated below: The weight of fuel divided by the weight of air in a given mixture is known a s the latter's fuel-air ratio (F/A). It will be noticed that the range of usable fuel-air ratio is relatively small.

1

MIXTURE LEAN TO BURN

,'&a

PS:I

2081

IS:/

-.-..

..-.en, OF I ---, .. .p OF AIR 8x1 ZSStl A/R FUEL RATIO (A/F)- WEIGHT OF A l e / WEIGHT OF FUEL

~ S : I

Fig. 4 -Combustible Range of Fuel and Air Mixture

PWA. 01. 100

PRAT'I' & WEIITNEY AIRCRAFT "BEST POWER" and "BEST ECONOMY" MIXTURES If the power output (brake horsepower) and rpm of an engine are kept constant, and the mixture strength is varied from one extreme to another, i t will be found that : 1. The manifold pressure required to produce the given power a t the given rpm will be a t a minimum when the fuel-air ratio is in the range of .074 - .080. By definition, .074 fuel-air ratio is "lean best power" mixture and -080 fuel-air ratio is "rich best power" mixture. Any mixture between ,074 and .080 fuel-air ratios is a "best power" mixture. 2. The quantity of fuel required to produce the given power a t the given rpm will be a t a minimum when the fuel-air ratio is approximately 0.0625 with the usual spark advance settings and compression ratios employed. This is, again by definition, the "best economy" mixture. The relation between best power and best economy are more precisely shown in Fig. 5. .IDEAL AND ACTUAL FUEL-AIR RATIOS Were there no considerations other than power and economy, the ideal fuel regulating device would be one which would give a fuelair ratio of 0.080 when maximum performance was desired, and which could be used to give a fuel-air ratio of 0.0625 when fuel economy was the primary consideration. Plotted in the form of a conventional carburetor performance curve, the relation between fuel-air ratio and power (sometimes expressed as airflow) are illustrated by the solid lines in Fig. 6.

IDLING

-

100 TAKE-OFF

50

PERCENT OF NORMAL RATED POWER

Fig. 6 -Ideal and Actual Fuel-Air Ratios

In practice, however, engine requirements call for deviations from this ideal curve, and the actual performance curve looks more like the one drawn in dotted lines in Fig. 6. The principal reasons for departing From the ideal fuel-air ratios are: 1. To insure smooth operation in the idling range, when the engine is cold. 2. To protect the engine from detonation in the high power range. .

IDLING MIXTURES

,@<;

.

.

''

---

rI - I

>

I 11

: II

Idling and taxiing present special conditions. . , The fuel-air ratio must be held within narrow limits : to prevent the spark plugs from fouling because of excessive richness, on the one hand, and, on the other, to avoid any tendency to - I"die" or accelerate hesitatingly because of over- ; leanness. Properly set idle mixtures will make ' ' L .. possible continuous smooth idling without . 1.1 danger of fouling. I 8

HELD

I

2.HOR5

FUEL AIR RATIO

Fig. 5 - Determination of Best Power and Best Economy

ixture ~rrengrn

8

.

..

::



CARBURETION

NORMAL COMBUSTION

will begin spontaneously and simultaneously throughout the unburned charge. The result is a sudden and violent explosion known as detonation.

DETONATION Normal combustion is rapid, but i t is by no means an instantaneous explosion. The charge burns evenly and smoothly, the flame front advancing a t a measurable rate-about 35 feet per second when combustion begins, increasing to roughly 150 feet per second, and finally slowing down a s the process nears completion.

Detonation occurs so quickly that even high speed cameras, which slow down normal cbmbustion to a snail's pace, fail to retard its progress sufficiently for exact analysis. It is accompanied by an abrupt pressure rise and violent pressure fluctuations of extreme rapidity. The engine is unable to turn into useful work energy so explosively released. The recurring shock pressures are carried to piston, cylinder, and hold-down studs, and the fatigue stresses set up

If sufficiently heated and compressed, any combustible mixture of gasoline vapor and air will catch fire. Accordingly, if the temperature and pressure of the unburned portion of the fuelair charge reach critical values, combustion

-

w

DETONATEO N

PWA. 01. 100 sure of the charge is too great, a critical value may be reached which will result in detonation. Excessive manifold pressure may be caused by too wide a throttle opening or, on some engines, by the use of too great a degree of supercharging. 2. Excessive carburetor air temperature (c.a.t).

- .,-.

As the temperature of the charge air a t the carburetor is increased, so is the temperature of the fuel-air mixture entering the cylinders. The latter is further raised during compression and combustion, and, ' if the initial temperature of the charge is too high, a critical value may be reached which will also result in detonation. Hot "free air" entering the induction system, inadequate intercooling in the case of multiple stage superchargers or too much carburetor pre-heat, may cause excessive c.a.t. High impeller speeds, the consequence of high engine rpm or of improper operation in the "high" impeller gear ratio, will cause a sharp heat rise through the supercharger, and, a s a result, the charge will not be sufficiently cool when delivered to the cylinders.

-

A few seconds of detonation can do this.

in the materials quickly lead to the failure of these parts. Detonation also causes a rapid rise in cylinder temperatwres, and thereby aggravates the very conditions which produced it. These high temperatures can rapidly destroy the piston, cylinder head, exhaust valve and guide, and damage other parts by burning and e>osion. Similar in its results to detonation, and frequently accompanied by it, is preignition. The latter is caused by uncontrolled ignition of the charge ahead of the normal flame front, because of contact with some "hot spot" in the combustion chamber, such as an incandescent spark plug. As a result the timing is too far advanced; the engine loses power and overheats ; local temperatures a t the hot spot rise rapidly; and the engine may be damaged, if it is not quickly stopped. Detonation-free operation is altogether normal and entirely possible over the full range of rated engine performance, even under the most adverse conditions. Nevertheless, detonation is the most likely a s we8 as the most destructive of the possible consequences of improper engine operation. CONDITIONS LEADING TO DETONATION Among the conditions which may lead to detonation the most important are :

1. Excessive manifold pressure. As manifold pressure is increased, so is the pressure of the charge entering the cylinders. The latter is multiplied many times during compression and combustion, and, if the initial pres-

.

3. Excessive cylinder head temperatures. The temperature and, indirectly, the pressure of the unburned portion of the charge may be raised to critical values a s a result of excessive cylinder head temperatures alone.

4. Improper grade of fuel. If the fuel used has an anti-knock rating (i.e., resistance to detonation) lower than that called for by the rating of the engine, detonation will follow any attempt to operate in the high power range.

5. Malfunctioning of the ignition system: Whenever the engine is operated in the high power range, detonation is likely to occur if the timing of the spark is too f a r advanced. I t may also occur during high power operation in a cylinder where only one of the two plugs is functioning. 6. Lean Mixtures. The tendency to detonate varies with the fuel-air ratio, and mixtures a t or near best power are the ones most likely to detonate. Combustion chamber temperatures may be lowered most effectively, and detonation thereby most readily inhibited by enriching the mixture beyond the best power setting.

-

-


THE AIRCRAFT POWERPLANT Detonation imposes one of the most important limitations on engine performance, and the operator must a t all times so control conditions a s to avoid any which might lead to detonation and the consequent damage to his powerplant.

.I 10

I

CARBURETION

RICHER MIXTURES-

. l o o mIN

THIS RANGE MIXTURE

P.OOO

5

a

EFFECT OF DETONATION

5.080 -I

W

As cylinder pressures are raised with an increase in power, the tendency to detonate naturally increases as well. Because of this it is necessary to deviate from the ideal fuel-air ratio, and gradually enrich the mixture power' is increased above approximate1 of Normal Rated power. The minimum fuel-air ratio necessary to protect the engine is determined by test, and the resultant departures from the ideal curve will look hmething like Fig. 7.

k

,070 UNECONWCAL T O OPERATE WtTH ECONOMY

L€A4E< -THAN 8EST .


Fig. 8-Complete

Fuel-Air Ratio Requirements

It is conceivable that mixture regulation could be accomplished by means of a manually operated valve inserted in the fuel line. But i t may be reasonably doubted that any pilot, or pilot's assistant, by continually taking stock of the conditions under which he is operating, and referring them to his fuel-air ratio curve, could

I

late the mixture with sufficient rapidity or

40

SO

60

70

--

--

% cy to ammmodate the constantly varying demands of the engine. Mixture regulation

.- -

PERCENT OF NOR'MAL RATED POW

-

Fig. 7 Departure From Ideal F uel-Air Ratio Required in Order to Suppress Detonation

THE COMPLETED FUEL-AIR CURVE

I ,

is accordingly turned over to a mechanical dev i a which performs this function automatically. Thia device is the carburetor. Bllaaa, airflow, or the weight of air consumed per hour, provides the link by which the fuel metming of the carburetor is coordinated pith the power demander of the engine. Mass airflow and power are direcay related, and, if the carburetor is made ~ensitiveto changes in airfiow, it can be made responsive to variations in p a m r

.

1-

After making the necessary adjustments for the idling and power ranges, the completed curve representing the engine's fuel-air ratio re- I quirements throughout its entire operating range will be somewhat like Fig. 8.


PWA. 01. 100

90

5t3

A-

P f f X E N T W N0.RMA.L Acl TED POW&?

*-

-

80 9b OR LBS/HR AIRFLOW

TAKE- OFF

Fig. 9 - F i n d Fwd-Air Curve

and so provide the fuel flow appropriate to each part of the engine's operating range. In other words, the fuel metering forces in the carburetor are adjusted to respond to airflow in such a way as to produce a fuel-air ratio power curve similar to Fig. 9, which is typical for all carburetors, and which closely resembles the fuel-air ratio curve just discussed. For purposes of discussion, carburetors may be divided into two general c l a s w : 1. Non-automatic. 2. Automatic.

through it, If altitude or temperature are increased, the same weight of air will occupy a greater vdume, and thus will flow through the carburetor with a higher velocity. This has the effect of increasing the fuel metering forces. The simple airflow measuring device of the non-automatic carburetor does not fully compensate for this extra force, with the result that the mixture h o m e s richer as temperature or altitude is increased. The effect on the basic metering curve af an increase in altitude is illustrated in Fig. 10. To compensate for this enrichment, a manually operated mixture control mechanism is ineorporated in the carburetor.

NON-AUTOMATIC CARBURETORS /"-')

As previously pointed out, proper fuel regulation depends on the ability of the carburetor to measure correctly the weight of air flowing

,"?I' I

/

I

/I

vMME T fRED F U f 1

I

By moving the mixture control lever, the area of the metering jet (orifice) is restricted or enlarged, and the fuel flow correspondingly decreased or increased. PERCENT O F N O R M A L RATED POWER OFf

Fig. 10 - Effects oF Altitude on Uncompensated Carburetor

If no adjustment is made, the fuel-air ratio will normally be somewhat richer than the minimum required by conditions in the cruising range. To realize the maximum fuel economy

/THE AIRCRAFT POWERPLANT possible for this type of operation, the mixture must be manually leaned. Non-automatic carburetors are usually restricted to the float type carburetors. AUTOMATIC CARBURETORS Automatic carburetors incorporate an automatic mixture control unit, sensitive to pressure and temperature, which compensates for changes in both altitude and temperature. These carburetors meter fuel in such a way as to conform to the basic fuel-air ratio curve throughout all flight conditions, and it is therefore unnecessary to vary the position of the mixture control lever to maintain the manually selected mixture setting. Two manually selected mixture settings are usually available : 1. Automatic Rich* (Take-Off .and Climb)

c .

2. Automatic Lean (Cruise)-to provide mixtures giving the greatest possible fuel economy; limited by engine power and grade of fuel ; and generally permissible only under favorable conditions of power and cooling.

The carburetor performance curve, with the two selective mixture settings, is illustrated by Fig. 11.


-

Fig. t l A Setting Curve for an Automatic Carburetor *The above mixture position designations have been used in conjunction with most engines manufacturedby Pratt & Whitney Aircraft and of other manufacture that are provided with pressure injection carburetors. Other designations may be used by other manufacturers. Recently some pressure injection carburetors have "rich" and "normal" mixture control positions instead of "automatic rich" ahd "automatic lean", and in the future different terms may apply.


CARBURETION

It is the usual practice to limit the use of the automatic lean setting to powers below 67% of Normal Rated power. More recently some automatic pressure injection carburetors have mixture positions as follows : Normal -This provides mixtures for all flight conditiops except in some instances for take-off and final approach. For low powers it provides a mixture strength comparable to the automatic lean setting of other carburetors. At higher powers the mixture is enriched to give mixture strengths comparable to automatic rich.

2.

Rich -This provides mixture strengths slightly richer than normal. I t is used for all ground operation, and when critical engine cooling requires a mixture richer than that obtained with normal. In many instances it is the required mixture control position for take-off and final approach.

Automatic carburetors may be of either the float or injection types. FUEL INJECTION. Once the fuel has been metered in proportion to airflow, it is introduced into the air stream as a fine, atomized spray. In some induction sys-, tems the fuel is atomized a t the carburetor; in others this is accomplished by a so-called spinner discharge nozzle which discharges the fuel into the impeller. In both systems the action of the supercharger further assists the process of atomization and vaporization. The above injection systems covered practically all aircraft engine applications prior to World War 11. An additional system which injects the fuel directly into each individual cylinder is being offered by several engine manufacturers as standard or alternate equipment. This system permits the control of the amount of fuel furnished to each individual cylinder and results in a more nearly perfect distribution of the liquid portion of the mixture. This advantage may be somewhat offset if there is lack of perfect air distribution and by the increased weight and complexity of this arrangement. Continued development should result in the more widespread use of the direct injection system. 31

r

0.

i

I

I

I

I I

I

I


PWA. 01. 100

CONTROLS

INSTRUMENTS

The functioning of the carburetor is regulated by two controls :

The fuel pressure gage is the only instrument connected directly to the carburetor. It is connected a t the fuel strainer chamber and the pressure registered on the gage is the differential pressure between that of the fuel and that of the atmosphere. (Generally, a t the carburetor entrance.)

1. Manual mixture control lever. This controls the fuel flow and so regulates the fuel-air ratio on non-automatic carburetors. On automatic carburetors it sets the mixture in either the Automatic Rich or Automatic Lean (Rich or rormal) position. (On some types of carburetors an additional position, known as Full Rich, is also provided. I t is basically slightly richer than automatic rich, and is not subject to compensation by the automatic mixture control units with the result that the mixture strength undergoes further normal altitude increase.) On both types of carburetors the fuel flow may be stopped by placing the mixture control lever in the Idle Cut-off position. This is the normal way of stopping an engine. (Idling mixtures, as well as idling speeds, may also be adjusted, but only on the ground.)

Throttle. As the carburetor is a t the air entrance to the engine, i t is a logical place a t which to provide some means of controlling the airflow. This is accomplished by the throttle, a valve which, by varying the area through which the charge air passes, controls the flow. Since airflow is directly related to power, the throttle is one, but not the only, means of controlling power. On pressure type carburetors the fuel metering force established in proportion to airflow is of sufficient magnitude to be used for regulating mechanisms whose functioning must vary with power. An example of this is the automatic spark advance mechanism.

FUEL PRESSURE The pressure of the fuel a t the entrance of the carburetor has an important effect on the fuelair ratio. If i t does not fall within the specified limits, the carburetor will not meter fuel correctly in response to airflow. The operator has no control of the fuel pressure produced by the engine-driven fuel pump while in flight. Consequently, he should check to see that it registers correctly during the ground tests and make all necessary adjustments prior to take-off.

In addition, three other instruments may be added to the induction system in order to assist in the control of induction system conditions and the regulation of fuel flow. These are : 1. Carburetor Air Temperature Gage :-The carburetor air temperature gage is usually attached in the air scoop elbow immediately upstream from the carburetor entrance, as the temperature measured a t this location is the most reliable means of determining ice formation conditions. It also provides the reading of gyeatest stability for determining the temperature condition of the air for the purpose of power calculation. The temperatures measured a t points in the induction system between the carburetor and the intake port do not provide a consistent means of measuring temperature and so do not provide a reliable means of determining charge density. By calibrating engine performance against variation in the temperature of the air measured a t the carburetor entrance the effect of the variations in temperature a t this point can be accurately determined. Attempts have been made to measure the condition of the charge air a t points between the carburetor and the intake port but they have not proved feasible, a s stable readings of a fuelair mixture cannot be obtained in the widely varying range of mixture strengths that are used.

i'

. -

-

2. Fuel Flow Meter: Fuel flow meters are sometimes provided on large aircraft in order to provide a close check on the rate of fuel consumed.

3. Fuel-Air Analyzer: On some installations, particularly those using non-automatic carburetoys, fuel-air analyzers are the only means of determining whether or not mixture strength is within the proper range. This instrument determines the fuel-air ratio by sampling the exhaust gases.

'<.

THE INDUCTION SYSTEM-2. SUPERCHARGING


2. SUPERCHARGING POWER AND MASS AIRFLOW The aircraft engine is a heat engine which derives its power from the burning of a mixture of fuel and air. The energy released by combustion is directly proportional to the weight, or mass, of the charge drawn into the cylinders. Accordingly, if the fuel-air ratio is kept constant a t the "best power" mixture setting, the power developed in the cylinders will vary directly with the mass of air consumed. This basic relationship is independent of rpm, except as the latter affects the engine's capacity to handle air.

driving the accessories ; while some of it is absorbed in driving the supercharging mechanism. What remains as useful power available a t the propeller shaft is known as brake horsepower, (bhp). In other words: ihp = bhp + power losses. Inasmuch as the power losses are functions of both rpm and power, bhp bears no fixed relation to rpm or to mass airflow. Although bhp is not strictly proportional to airflow, the connection between the two is nevertheless close, and in the, final analysis the power output of an engine is cbntrolled by the mass or weight of air flowing through the carburetor in a given time. To obtain a hidh power output, it follows that the designer must provide some means of supplying the engine with a mass airflow sufficient for all conditions where high performance is demanded. To do this satisfactorily over a wide range of altitude, and without too great a rise in the temperature of the fuel-air charge, or too great an increase in the complexity, weight, and drag of the installation-poses one of the most difficult problems in aircraft engine design. -

-

MASS AIRFLOW THROUGH CARBURETOR (LB PER HR 1

Fig t 2. - Relation Between Power (IHP) and Mass Airflow

The power developed in the cylinders of an engine is known as the indicated horsepower (ihp). Not all of it is delivered to the propeller shaft, but, as suggested below, some of the ihp is lost in overcoming internal friction and in

THE ENGINE AS AN AIR PUMP Air, or more strictly speaking, the oxygen in the air, is by volume and by weight the most important part of the combustible mixture. Roughly 13 pounds of air, or 2.6 pounds of oxygen, are consumed in the combustion of 1 pound of gasoline a t "best power" mixture strength.

-

PWA. 01. 100

PRA'IT & WHITNEY AIRCRAFT .

Unlike the rocket, which to be self-contained must carry both ingredients, the conventional aircraft carries only one ingredient in its tanks and flies through the other.

It is an absolute pressure, measured from zero

The internal combustion engine may, accordingly, be thought of as an air pump, whose performance depends in large measure on its capacity. Half of each power cycle is devoted to pumping air. As the piston descends during the intake stroke, air is drawn into the cylinder. During the exhaust stroke it is returned to the atmosphere, as a part of the exhaust gases, by the pressure of the piston. In the course of the compression and power strokes the engine takes time out from pumping to produce enough power to keep itself going and deliver a reasonable balance a t the propeller shaft.

Any variation or restriction of airflow affects the pressure a t the collector rim. Manifold pressure is, therefore, regulated chiefly by the carburetor throttle valve. Secondarily, it is controlled by the pressure which exists a t the car-

psi, and not a "gage", or differential pressure, measured from some reference point, such as atmospheric pressure.

TAKE P/PE

MANIFOLD PRESSURE

buretor entrance a s the result of altitude, ram, or the operation of an auxiliary air compressor or supercharger. .

.

The weight of air consumed by an engine depends primarily on the total piston displacement, the rpm, the temperature of the charge, and the pressure existing a t the intake ports. If the last of these factors is increased, a greater quantity of air is forced into the cylinders, and, so long a s the fuel-air ratio and other factors remain constant, the result is to increase the weight of the charge. Since the energy released by combustion is in turn proportional to the weight of the charge, it follows that intake port pressure is an important index of power whenever the other factors are known.

,

In practice this pressure is not measured at an intake port, but a t the rim of the supercharger collector. It is known as manifold pressure, or absolute blower rim pressure (abrp), or even manifold absolute pressure (map), and is expressed in inches of mercury (in. Hg.).*

* In. Hg. is the measure used in the United States of America. Other countries ma use different units; for example in England manifollpressure is expressed in psi "boost" above sea level atmospheric pressure.

Manifold pressure, it should be noted, is not power, but merely a convenient index for measuring one of the several factors affecting power. It is only in combination with rpm, and after making certain necessary corrections for carburetor air temperature, that manifold pressure may be used in making an approximate determination of bhp by reference to the appropriate operating curves.

CARBURETOR AIR TEMPERATURE AND MASS AIRFLOW Any factor which changes the mass of the fuel-air charge will affect the engine's power output. The effects of variations in manifold pressure have been noted. Another important factor is the temperature of the charge. Since the weight of a given volume of the fuel-air mixture a t a given pressure is inversely proportional to its absolute tempwature (abs. temp. = F 459 or C 273), it follows that any increase in the temperature of the charge a t a given manifold pressure results in a decqease in

+

+



SEA n

V

L

L

IUVVV

PUUV '

Fig. 73

-----

ALTITUDE FEET

- Power vs Altitude - Unsupercharged Engine

its density and weight, and hence in the amount of energy available from its combustion.

'

SUPERCHARGING

In practice the temperature of the fuel-air charge itself is not measured. Instead, the tem- ' perature of the charge air is taken as it is about to enter the carburetor prior to the injection of the fuel ;and, in establishing safe carburetor air temperature limits, allowance is made for the cooling effect of fuel vaporization, for the heat rise due to compression in the supercharger, and for other temperature changes occurring between the carburetor and the intake pipe. -

the wide open position, manifold pressure will be found to decrease as altitude increases, and the impression of a loss of power would be confirmed by a torquemeter. The reason for this is to be found in the ability - or, better, inability - of the engine to pump sufficient air.

High carburetor air temperatures are undesirable, not only because they result in a loss of power, but also because they may lead to detonation and consequent engine failure. It thus becomes one of the principal functions of the induction system to deliver the fuel-air charge to the cylinders a t the proper temperatures.

THE UNSUPERCHARGED ENGINE If the fuel-air charge is forced into the cylinders simply by the pressure exerted by the atmosphere, the engine is called a "naturally aspirated" or unsupercharged engine. ("Aspirate" means to draw by suction.) Such an induction system satisfies the requirements of moderate performance a t low altitudes. Primary training planes are often equipped with such engines.

Assume, for example, that an engine delivers 500 bhp a t sea level and, to produce this power, consumes in one hour 43,000 cu f t of air weighing 3,300 ib a t sea level. At an altitude of 10,000 ft, 43,000 cu f t air weighs 2,460 Ib and a t 19,000 f t weighs only 1,810 lb. Since the power output of the engine depends on its ability to consume pounds rather than cubic feet of air, it follows that the power will decrease with altitude much as shown in Fig. 13.

However, once this type of engine has left the ground its limitations become apparent. If the rpm is held constant and the throttle fixed in

The relation between altitude, weight, and volume can be illustrated in another way. At sea level, the 3,300 Ib of air occupy 43,000 cu f t

PWA. 01. 100

PRATL' & WHITNEY AIRCRAFT

- about' the size of a two-story building.

INCREASING THE PUMPING CAPACITY

At

The problem of designing an engine that will maintain its power output a t altitude thus resolves itself into one of increasing the engine's air-pumping capacity. Three different methods may be used to accomplish this :

1. Increase piston displacement by adding or enlarging cylinders. This is a cumbersome method, since i t increases the size of the engine without increasing. the ratio of power to weight. 2. Increase rpm. This raises the pumping rate of the engine, but offers only a partial solution to the problem, as centrifugal and reciprocating forces shortly impose a limiting speed.

3. Supercharge by incorporating in the engine a special air pump whose primary function is t o compress more weight of air into the fixed volume of air handled by the pistons. This is the method most commonly used to raise an engine's air pumping capacity, since it calls for no increase in rpm or of power section size, and adds relatively little to the overall weight of the powerplant.

10,000 f t this same weight will occupy 58,200 which will require an extra story on the cu ft

-

THE SUPERCHARGER

-

building. At 19,000 f t i t will occupy 78,000 cu f t - and the original two-story building is now three and one-half stories high.

The conventional supercharger is a centrifugal air compressor placed between the carburetor and the intake pipes. It is usually housed between the power and accessory rear sections. Its principal features, which are shown on p. 37, consist of three main units: 1. Impeller. After leaving the carburetor, the air passes through the supercharger throat t o the impeller. The impeller is driven a t roughly 6 to 14 times crankshaft speed, and because of its high rotational speed imparts a large velocit y energy to the air. 2. Diffuser. As the fuel-air charge leaves the impeller i t passes into the diffuser. The vanes of the diffuser ensure a smooth flow while allowing the charge to slow down as i t moves outward, with the result that the velocity pressure acquired from the impeller is transformed into static pressure.

3. Collector. After leaving the diffuser, the charge is stored momentarily and equalized in the collector, whence i t is drawn to the cylinders through the intake pipes. I t is a t the collector rim that manifold pressure is taken.



BUPERCHARGING

TH/S /S KNOWN AS A CENTR/FUGAL COMPRESSOR

CARBURETOR

+----

GEAR TRAIN

CRANKSHAFT

DIFFUSER VANE

COLLECTOR

INTAKE PIPE

SECONDARY EFFECTS OF SUPERCHARGING While the princi~alfunction of supercharging is to increase mass airflow by raising the pressure, and hence the density and weight of the fuel-air charge, several other important effects are also associated with it. These are: 1. More even distribution of the fuel-air charge to the cylinders, because of the radial construction of the supercharger and its axial location.

2. More complete vaporization of the fuel. The fuel is injected into the air stream a t the carburetor or a t the entrance to the supercharger through a spinner discharge nozzle attached to the impeller shaft, Vaporization is assisted by the whirling action of the impeller a s well a s by the heat of compression.

3. Rise in temperature of the fuel-air charge as a result of compression. While this assists vaporization, a t the same time i t tends to lower the density of the charge and so reduce power a t any given manifold pressure. In extreme cases these high temperatures may lead to detonation. The temperature rise increases rapidly with impeller speed, and, as a consequence, an engine may actually be able to deliver more power a t a low rpm than a t a high rpm, since the lower charge temperatures permit operation a t higher manifold pressures without danger of detonation.

4. Absorption of power by the supercharging mechanism. Additional pumping power is necessarily required for any increase in airflow. Since the power required to deliver a given weight of air varies roughly a s the square of the tip speed of the impeller, it follows that, for any given

PWA. 01. 100

PRA'IT & WHITNEY AIRCRAFT ipm and manifold pressure, the power absorbed by the supercharger will be kept t o a minimum by operating in the lowest possible impeller gear ratio that will provide the desired engine performance. "GROUND BOOSTING" Suppose a supercharger is installed in the induction system of the normally aspirated engine previously used as an example. The cylinders will still pump 43,000 cu f t of air per hour, but the supercharger will compress a greater amount of air into this volume. Assume that the equivalent of 60,000 cu f t of outside air are compressed into the 43,000 cu f t per hour capacity of the cylinders. At sea level 60,000 cu f t of air weigh 4,600 lb, while 43,000 cu ft, a s previously noted, weigh only 3,300 lb. Since indicated horsepower (ihp) is proportional to mass airflow, the engine should respond to this increase in airflow by producing 700 horsepower in the cylinders, whereas, with the original airflow it produced only 500. Neglecting friction and other power losses, 50 horsepower must now be diverted to drive the impeller, so the net gain in available brake horsepower (bhp) is 150. JOOIW

1 I

1

E

3 U

i

UNSUPERCHARGED

SUPERCHARGED

The performan&-of the supercharged engine compared with that of the normally aspirated engine is illustrated in Fig. 14.

Fig. 74

3

8

If the engine is built sturdily enough to absorb the additional loads imposed by the higher cylinder pressures which result from supercharging, and if the fuel does not detonate because of increased pressures and temperatures, the engine may be run a t sea level with its throttle wide open, and its full power capacity may be called upon a t any altitude. Such a supercharged engine is known as a "ground boosted" engine, and is typical of the power plants of many basic training airplanes. ALTITUDE SUPERCHARGING The air pumping capacity of the engine may be increased by enlarging the size of the impeller or by driving i t at a higher speed relative to the crankshaft. The degree of supercharging obtainable by either of these methods is, however, subject t o two limitations: 1. Temperature Rise Through the Supercharger : The compression of the fuel-air charge is accompanied by a temperature rise attributable in part to compression and, in part, t o fluid friction and turbulence. It is desirable to keep this rise to a minimum for the purpose of obtaining maximum charge density and, more important, the amount of this rise . determines the limitations of power which must be imposed to prevent detonation because of excessive charge temperatures. This last consideration is the most important limitation placed on the tip speed a t which the impeller may be driven.

-Power vs Altitude - Unsupenharged and Supercharged Engines


THE INDUCTION SYSTEM-2. -

SUPERCHARGING

To obtain increased performance from the engine previously used as an example, a higher capacity supercharger is installed in place of the one which provided the "ground boost." The impeller of the new supercharger has a higher tip speed, and is assumed to operate just within the detonation limits imposed by the temperature rise. It supplies the cylinder pumping capacity of 43,000 cu f t per hour with the equivalent of 85,000 cu f t of air per hour. At sea level 85,000 cu f t of air will weigh 6,500 Ib, and the power potentially available a t the crankshaft (again neglecting friction losses) will be increased from 700 to 1,000 horsepower. The higher capacity impeller, however, absorbs 100

Power Absorbed by the Supercharger: As previously stated, the power required to deliver a given weight of air varies approximately as the square of the impeller's tip speed. Conceivably, an impeller speed could be reached above which i t would no longer be efficient to drive the impeller because of the excessive fraction of the engine power absorbed by the supercharger. However, the limitations imposed by temperature rise will be reached considerably in advance of this speed.

horsqwwe~,Instead of 50, leaving 900 p o h tially a'traikble at the propeller shaft.

The performanes of the ensins equipped with the h i g k capacity supercharger would be illusOrated by the chart in Fig. 15, provided it were possible to w e at every altitude all of the power Ulm potentially avrilabb.

ALTITUDE FEET

Fig. I S -Power vs Altitude

- Highly Supercharged Engine

PWA. 01. 100


5000

10000

I5000

20000

25000

ALTITUDE FEET

- 87 Octane

Fig. 16 -Descent Performance

SUPERCHARGING LIMITATIONS

Still more power is potentially available at a lower altitude, but a t the given maximum permissible rpm this additional power is not attainable without detonation. Below 13,000 f t the engine must accordingly be operated a t part throttle, and the maximum bhp t h a t can actually be used is 570. This power is available a t part throttle from 13,000 f t , the so-called full throttle, or critical altitude t o sea level; above 13,000 altitude the 570 bhp cannot be obtained.

A maximum of 900 bhp is only potentially available, however, a t the propeller shaft. How much of this potential power can actually be used may be determined, if the engine is taken to altitude, and a very gradual descent begun from 25,000 f t a t full throttle and maximum permissible rpm. The descent is commenced , using 87 octane fuel-i.e., a fuel of relatively 1ow::I anti-knock value. As altitude is decreased thz; pressure and temperature of the charge air in-. crease, until a t 13,000 f t they reach t h e critical values for this particular type of fuel, and' detonation begins. A t this point the engine is developing 570 bhp.

W-f!

SEA LLVEL

5000

The fuel selector valve is now switched t o a tank containing 100 octane fuel-i.e., one of high anti-knock value -and the descent continued from 13,000 f t a t maximum permissible rpm and full throttle. A t 10,000 ft, corresponding t o a

10000

I5000

ALTITUDE FEET

Fig. 17 - Descent Performance

- 100 Octane

20000

25000



bhp of 635, the pressure and temperature of the charge become critical for the 100 octane fuel, and detonation again begins. With this fuel the power output must be limited to 635 bhp, and is available a t part throttle, from sea level to 10,000 ft, the critical altitude.

SUPERCHARGING

COMPARISON OF NATURALLY ASPIRATED, GROUND-BOOSTED, AND ALTITUDE ENGINES The difference between a ground-boosted and an altitude engine is largely a matter of definition. For both normally aspirated and groundboosted engines the maximum usable power and the engines' theoretical potential power produo ing capacity are the same. In the case of an altitude engine, on the other hand, the maximum usable power is less than the engine's theoretica1 capacity. It is also conceivable that an engine might be considered a ground-boosted engine with a high quality fuel and an altitude engine with a lower grade fuel.

The selector valve is now switched to a third tank containing a magic, non-detonating fuel, and the descent is continued as before. With this fuel, detonation is no longer the limiting factor. Nevertheless, a t 4,000 f t there is an ugly noise and the engine fails as it is developing 790 bhp, indicating that the limit of its structural strength has been reached. This, then, is the maximum available bhp regardless of fuel quality, and 4,000 f t is the correspondir~gfull throttle, or critical, altitude, When the engine failed a t 790 bhp, an additional 90 horsepower were being absorbed by the supercharger. (Friction and other power losses are neglected.) The total of 880 horsepower is thus the ihp capacity of the engine as determined by its structural strength regardless of the degree of supercharging. The engine just described can therefore never develop the 1,000 ihp or the 900 bhp potentially available a t sea level ; and all operation from sea level to the critical altitude--as determined by the fuel or the engine's structural limitationamust be a t part throttle. Such an engine for which full throttle operation a t sea level is restricted or prohibited, is known a s an "altitude" engine.

The foregoing discussion dealt with three variations of the same basic engine. The power sections and rpm were assumed to be identical in each case, the only difference being in the degree of supercharging. The performance of these three engines is compared in Fig. 19. I t will be noted that the ground-boosted engine, because so little of its power is directed to the supercharger, delivers more power (bhp) to the propeller shaft a t sea level than the highly supercharged engine using either 87 or 100 octane fuel. The better low level performance of the ground-boosted engine is more than offset, however, by the greatly superior performance of the highly supercharged engine a t altitude.

C

3 E A LEVEL

,

Fig. 18

10000

15000

ALTITUDE FCgJ

- Desced Pe-docl~~~nee - Nm-Daboncrting F w d

254300

i

R

Fig. 19

- Performance Comparison - Unsupercharged, Ground Boosted and Altitude Supercharged Engines

The restriction placed on maximum power outu t by structural limitations should not be ed as a reflection on the engine. The s problem is to maintain some desired rformance up to the highest possible altitude t h the least weight and size, and with the available quality fuel. This is achieved on the supercharged, or altitude, engine by proa pumping capacity t h a t is greater than

necessary or safe to use a t or near sea level in order t o meet the greater pumping requirements a t the higher altitudes. Since the danger exists t h a t the operator may inadvertently make use of this excess power a t low altitudes by exceeding a safe limiting manifold pressure, i t is essential that he exercise both care and judgment to protect the engine against the results of unrestricted operation.

3. DUCTING e portion of the induction system t h a t is of, or "UP-stream" from, the carburetor ally furnished by the aircraft manufacurer (as distinct from the engine manufac). While not part of the engine bill of mas, this ducting is a n intimate portion of he induction system and exerts considerable fluence on engine performance and the mainenance of proper operating conditions. The aircraft ducting affects engine operation

Because of the forward motion of the aircraft onsiderable velocity energy is available in the ir entering the carburetor air intake. Proper ormation of the air entrance and the passages nd elbows of the air duct can result in the reovery of a large portion of this energy, or am, and air will be furnished to the carburetor a t a pressure greater than t h a t of t h e altitude

a t which the aircraft is flying. This has t h e effect of increasing the supercharging capacity and, in extreme instances, high velocity aircraft can recover ram equivalent t o an additional stage of supercharging.

INDUCTION SYSTEM ICE PREVENTION AND REMOVAL To protect the engine induction system against ice formation, or to remove ice already formed, the ducting to the carburetor is arranged so that the ram air can be closed off and heated, or moisture-free air drawn in from a different source. The degree of heat required differs between engine models because of differences in the method of injecting t h e fuel. The greatest preheat capacity must be furnished for installations of engines in which the fuel is injected immediately downstream from the carburetor. In these cases t h e fuel evapora-" tion occurs in a region in which no heat is re-!:

THE AIRCRAFT POWERPLANT /'

-

..

FIDRWAL RAM,ENTRANCE7

THE INDUCTION SYSTEM-3. HOT AIR DOOR 7

N
-

/

EXHAUST COLLFCTOn

RLZERED AIR ENTRANCE

DUCTING

7

.

/ AIR /S WARAWE0 BY PASSING OVER THE E X M U COLLECTOR AND IS ADAWITTED I N ;W THE /NOUCT/ON J.--.-I-PA,J I C M n, HEN REOUIIQt-0 FOR IC, QEMOVAL OR PREVEN TIION

,..

,

.

ceived from other sources and the temperature of the charge is lowered as much as 50 F (28 C) with the result that the moisture contained in. the charge will freeze on any convenient surface and possibly block the passage if the carburetor air temperature is between 32 F-80 F (0 C-25C). When the fuel is injected immediately adjacent to the face of the impeller, the mechanical action of the rotating impeller practically eliminates this possibility and i t is not necessary to use the quantity of preheat for ice prevention that is required by other injection systems. However, regardless of the type of fuel i jection used, ice can be formed because of near freezing moisture in the atmosphere or because of the temperature drop past the throttle. All experience accumulated to date conclusively demonstrates that ample preheat capacity is the only sure method of providing the temperature rise needed to remove ice already formed. The so-called "alternate air systemsJ' taking air from over the cylinders or from the accessory compartment does not provide sufficient temperature rise to give this essential protection.

DUST PROTECTION Dust protection was not seriously considered until World War 11, but large-scale operations a t training bases in the United States and under combat conditions in North Africa and other theaters clearly demonstrated that when i t was necessary to operate in dust-charged air, protection for the engine must be provided. While dust to most individuals is merely an

annoyance, it is a serious source of trouble to an engine. Dust consists of small particles of hard abrasive material. Taken in by the air, i t can collect on the metering elements of the carburetor, upsetting the proper relation of ' fuel flow to power. It next asserts itself in the cylinder walls by grinding down these surfaces and the piston rings. It then contaminates the oil and is carried through the engine causing the bearings and gears to experience distress. In extreme cases an accumulation may clog an oil passage and cause oil starvation. Whereas dust is most evident close to the ground, in certain parts of the world i t can be carried to altitudes as high as 15,000 feet. Continued operation under such conditions withclilt engine protection will result in evidence of engine wear a s indicated by high oil consumption. During World War I1 in the North African operations, engines which had demonstrated the ability to perform satisfactorily for 600 or 700 hours between overhauls under dust-free combat conditions, could swallow dust for only 20 hours before requiring removal. When operation in dusty atmosphere is necessary, engine protection can be obtained by means of suitably designed alternate air entrances to the induction system which incorporate a dust filter. The system shown on the title illustration is typical. As the air entrance to the filter does not face directly into the air stream, considerable dust removal results in forcing the air to turn a s it enters the duct. The dust particles, being solid, tend to continue in a straight line and the greatest portion of them are removed a t this point. Those that are drawn in are quite easily removed by the filter.

I I

PWA. 01. 100

PRATl' & WHITNEY AIRCRAFT The efficiency of tkie filter depends upon proper maintenance and servicing. Periodic removal and cleaning of the filtering elements is essential to satisfactory engine protection. Most production aircraft for civilian use do not include filters a s standard equipment. Where the operations are conducted from wellestablished dust-free bases, the omission of the filtering system is justified in order to eliminate the appreciable weight and complication of the installation. In those cases where the aircraft is furnished with a filter and i t is possible to select either ram or filtered air, the ram position is preferred for all dust-free conditions of operation a s the filtering system causes a measurabIe loss of ram energy in the a i r flowing t o the engine.

SCREEN Pratt & Whitney Aircraft furnishes, with each engine, a coarse mesh screen to be in-

,

stalled between the carburetor and the adjacent portion of the aircraft charge-air duct. This screen is for the essential purpose of preventing stones, nuts, bolts, papers, rags, tools or any other loose object from further penetrating the induction system and damaging the impeller. Many operators have removed these screens in order to dispose of a possible surface for ice formation. While they were successful from this point of view their experience has been an unhappy one in that the inevitable stone picked up by a propeller, or the nut that was lying loose when the air scoop was removed, chewed up the impeller and complete engine removal was necessary. If the operating personnel can exercise the necessary precautions t o prevent ice (a factor over which they have effective control) i t will not be necessary to remove this protection t o the engine from conditions over which they have little, if any, control. Pratt & Whitney Aircraft strongly recommends that these screens be used a s intended.


THE IGNITION SYSTEM

THE IGNITION SYSTEM

p",po';x 3

HARNESS

SPARK PLUGS

1 8. Ignition Switch provides a means of turn:

and their functions are: 1. Magneto :generates a current of high voltage which is discharged a t the proper time. . . 2. Distributor: distributes the current to the proper spark plug. 3. Harness: transmits the current from the distributor to the plugs. 4. Spark Plugs: ignite the fuel-air charge in the cylinders. 5. Shielding : provides a grounded metallic enclosure for the entire ignition system to protect radio against interference from the ignition system. 6. Venting System: provides natural air motion or produces positive air flow to prevent contamination of the ignition system by moisture, engine oil or acids formed by electrical action. 7. Booster or Induction Vibrator: provides an auxiliary spark a t starting only.

ing the ignition either on or off. For safety reasons two ignition systems are normally furnished. In the event of failure of one, the other will insure continued operation. At the same time dual ignition, as this arrangement is called, will help the combustion process by providing two flame fronts which results in more efficient burning of the charge. CONTROL AND CHECKS Engine performance furnishes the only d u e to the functioning of the ignition system, and the pilot has no control over it other than to turn i t off or to turn on either or both halves of the system. A-

PLUGS FIRING

SWITCH POSITION

Front and Rear Rear (in each cyl) Front (in each cyl) None

Both

Left (L) Right ( R )

Off

PRATT & WHITNEY AIRCRAFT By operating the engine a t reduced power for a few seconds, first on one magneto, and then on the other, and noting its rpm with the ignition switch in the BOTH, LEFT, RIGHT and O F F positions, it may be possible to learn something about both the nature and location of an observed or suspected ignition irregularity.

IRREGULARITIES AND FAILURES The ignition system furnished with an engine is the result of careful design and thorough proof testing. In all probability it will give dependable service provided it is properly maintained and serviced. However, irregularities can occur which will affect its performance and while these difficulties are rare, it is well t o review their causes in order that the symptoms of maIfunctioning can be understood. 1. Breakdown of insulating materials, burning of breaker points, short circuits or broken electrical connections are comparatively rare. They are revealed by rough running, loss of power, or possibly complete stoppage.

More common are irregularities due to moisture which might accumulate in different parts of the ignition system. This accumulation can be caused by breaks, or loose covers allowing direct access of water to the system. Breathing, which occurs during the readjustment of the system from low to high atmospheric pressure, can result in drawing in moisture-laden air. Ordinary condensation of moist air, as the engine cools, can also result in appreciable moisture accumulation. This moisture causes the insulating materials to lose their electrical resistance. A slight amount of moisture contamination will cause a reduction in magneto output by leaking part of the current, intended for the spark plug, to ground. This condition will first be noticed in starting difficulty, a t which time the magneto output is already reduced by the low cranking speeds. If the moisture accumulation is serious, the entire magneto output will dissipate to ground and might cause flashover and burning of insulating parts, which can result in their permanent failure.

PWA. 01. 100 Moisture accumulation in flight is extremely rare, as the temperature of the system is effective in preventing condensation, and difficulties from this cause will more probably be evident on the ground.

3. The normal unpressurized ignition system can operate satisfactorily a s high a s 25,000 feet. Above this altitude, the low electrical resistance of the reduced air density will quickly show up any irregularity that would not be serious a t lower levels. As altitude is further increased, it becomes necessary to supply the system with air under pressure or to fill it completely with a nonconducting substance, in order to maintain sufficient electrical resistance to prevent flashover. 4. Probably the most commonly blamed of all causes of irregular engine operation is the spark plug.

This may be the result of incorrectly selected plugs, incorrect gap clearance, faulty installation, improper engine operation, carbon or lead fouling, condensation in the cylinders, or actual breakage of the plugs. On the one hand, the plugs may run so hot a s to become incandescent and cause pre-ignition and tend to lead fouling; on the other, they may run so cold as to be fouled out by the accumulation of unburned carbon. Spark plug trouble is likely to be localized a t one or several plugs, but the evidences of spark plug trouble are much the same a s those of other types of ignition irregularities. The cutting out of one-haIf of the ignition system, or the failure of a spark plug in one or more cylinders, will result in slower combustion in the cylinders affected. The effect is comparable to that of delaying the timing of the spark discharge, and the consequence is a loss of power, although rpm and manifold pressure will, a s a rule, remain unchanged. If both plugs fail in one or more cylinders, the result is, essentially, an engine of fewer cylinders. In any case, the attempt to recover the lost power by increasing manifold pressure is unwise, since it may lead to detonation, especially when operating in the high power range.

PART 11 POWER, BMEP, AND RATINGS Since engine design and operation both have as their ultimate objective the efficient and reliable production of power, it may not be altogether inappropriate a t this point to inject a definition of "power," together with a brief description of the methods by which it is measured and produced, and to conduct a hasty investigation of certain of the Iimitations to which its application is subject. In the course of this discussion the reader must desert the pilot's cockpit and the mechanic's, hangar for the laboratory of the engineer and the ivory tower of the physicist'. If in these

new surroundings he encounters a few faces, or if the ill-laid ghosts of some forgotten course in elementary mecha t r y his patience and shatter his in indulgence is craved; and it is to be ho his industrious application to the du cult paragraphs that folIow in PART I1 will rewarded by a t least a nodding acquainta with such formidable terms a s indicated horsepower, brake mean effective pressure, and Nor ma1 Rated power, a s well as by a be standing of some of the problems fu to powerplant design and operation.

POWER AND ITS MEASUREMENT WORK (Force x Distance)

OR nition of power involves the three force, distance, and time. The ime and distance are deceptively do not permit a ready definition, anings will none the less be genervlous. Force masquerades in a variety of es - mechanical, electrical, chemical ntially i t is anything which tends to ibit, or alter motion, such as a presht, a tension. Motion is not necesplied, since two or more forces-acting eously on a body may be so balanced as r the body immobile.

To many individuals work is the effort e pended (if any) between "punching the clock" in the morning and "ringing out" a t night, and is measured in terms of fatigue, boredom, or dollars and cents. The engineer's definition, while lacks human interest, is somewhat more precis For him, work is the product of a force by the distance through which i t acts, or: Work = Force x Distance (in the direction of the force) For example: If a weight of 50 lb is raised a distance of 11 ft, the work done is: 50 Ib x 11 f t = 550 f t lb. The same amount of work, 550 ft-lb, will done if a 10 lb weight is raised 55 ft.

but to the initiated i t is the rate of doing

Power = Work/ Time If the reader climbs 4 flights of stairs, the work done -i.e., the product of his weight by the total height of the stairs - will be the same regardless of the rate of ascent. On the other hand, if the reader first saunters slowly up the 4 flights and a little later, being sound in wind and limb, runs up them in, say, one-third of his original climbing time, pulse and perspiration should convince him that he develops considerably more power during the rapid ascent than during the leisurely climb.

Power is expressed in a number of units horsepower, kilowatts, Btu per minute, chevalvapeur, etc. - of which the first named is perhaps the most familiar. One horsepower (hp) is defined as work done a t the rate of 550 ft-lb per second, or, what is the same thing, 33,000 ft-lb per minute. No. ft-lb per sec 550

-

No. ft-lb per min 33.000

: if an engine is doing work at the rate of

0 ft-lb per sec, or 990,000 ft-lb per min, i t is oping 30 hp.

It would be theoretically possible to measure the power delivered by an aircraft engine a t the propeller shaft - i.e., the engine's brake horsepower (bhp)-by a method similar to the one illustrated on page 49, if the propeller shaft were made into a drum and substituted for the hand operated drum in the sketch. Such a method, however, presents obvious difficulties, and in practice the measurement of an engine's bhp involves the measurement of a quantity known a s torque, or twisting moment. An unpleasant form of torque is the twisting force that must be applied to the crank of a recalcitrant outboard motor. More precisely, torque is defined as the product of a force by the distance of the force from the axis about which i t acts; or: Torque = Force

x Distance ( a t right angles to the force)

For example: if a force of 20 lb is applied a t the end of a 3-ft-arm, a torque, or twisting moment, of: 20 lb x 3 f t = 60 lb-ft is exerted on the shaft to which the arm is attached. I t should be noted that shaft and arm need not be rigidly atached to one another, and that the torque exerted on the shaft is independent of the latter's 2 SECONDS

i

a s a 20 Ib force is applied a t the

exerted by a propeller shaft can exerts a torque of 80 Ib-ft, i t

horsepower delivered may be read of i t s So lor the sl or no1 .

THE PRONY BRAKE There are a number of devices for meas torque, of which the Prony Brake, dynamomet and torquemeter are examples. The Pro

is properly, though not consistently, expressed in Ib-ft; and is not to be confused with work,

described briefly on page 50, may be t a typical of these devices.

The Prony Brake consists essentially of a collar, or brake, which can be clamped drum splined to the propeller shaft, and arm of known length terminating a t

ever, is prevented by the arm, and the ssary to arrest the motion of the arm

puted without difficulty. Assume, to begin with, that the engine's propeller shaft is stationary while some obliging chap rotates about i t the entire Prony Brake mechanism, scales and all. This will require work, since the friction between the propeller shaft drum and the brake collar must be overcome. The force appIied a t the end of the arm to balance-i.e., just overcome-the resisting friction is shown on the scales. The distance through which the man must push this force in the course of one revolution is: 2~ x Length of Arm. The product of the force by the distance through which i t is moved (or acts) is thus the work performed by the obliging individual ; or :

ating shaft. Hence :

Work per Rev. = Force (Weight on Scales) x 27r x Length of Arm and since: Torque = Weight on Scales x Length of Arm Work per Rev. = 2w x Torque

ple: if the scales register 200 Ib and the the arm is 3.18 ft, the torque exerted by 200 Ib X 5.18 f t = 636 lb-ft

REVOLUTION ( 2 x~Torque) the propeller shaft can be com-

Since the friction arises from the relative motion of propeller shaft drum and brake collar, the same amount of work will be done by the propeller shaft in one revolution a s was done by the man, if the shaft turns and the Prony Brake meck~anismremains stationary.

-

FRICTION ADJUSTING WHEEL

SCALES 200 LBS.

PROPELLER SHAFT

LENGTH OF ARM

cales register 200 Ib and the before, the work done by the revolution is: 200 lb X 2n X Or, if i t were known that the orque of, say, 1000 Ib-ft, tke work n of the shaft would be: 2?r x 1000

(K x Torque x rpm) volution is multiplied by the e the work per minute, or per minute is expressed in and this quantity is divided by t will be the horsepower output k per Revolution x rpm Torque) x rpm

ind therefore :

,000

= 1.904 x lo-' if torque is

the scales read 200 lb, the arm is 3.18 m of the shaft is 1650, the horse1650 rpm/33,000 = 200 hp. Or, if the torque of a shaft rotating a t 2000 rpm is kno\r,n to be 330 lb-ft, the horsepower being delivered i s : 1.904 x lo-' x 330 lb-ft X 2000 rpm = 1260 hp. The horsepower course, the engine's brake horsepow

IRQUE -RPM g as the fricti rnd propeller shaft dru

pose an appreciable load on the e so great a load as to prevent the en turning, it is not necessary to k amount of friction between collar no load were imposed, there would to measure and the engine would the load were increased to stopped the engine, while there to measure, there would be no ease would i t be possible to gine's bhp.

down because of the greater load imposed, with i t becomes greater, the arm bears heavily on the scales, and the torque So long a s the increase in torque is to the decrease in rpm, the h by the shaft remains unchanged-a out by the equation: This is an important relationshi that horsepower is a function of bo rprn. For example: and engine wi1 same horsepower a t the propeller each of the following conditions :

PWA. 01. 100

WHITNEY AIRCRAFT the friction b as an aircraft engine is running a t a ed, the torque exerted by the enne a t the propeller shaft must equal the opposg torque exerted by a Prony Brake, propeller, other load-imposing device. If this balance is ed-as by changing the throttle setting propeller blade angle - the unbalanced or decelerate the engine equilibrium is restored, or until the ails because of overspeeding or is stop. Since the rpm is necessarily both propeller and propeller shaft, t the power absorbed by the pror a t constant rpm equals the power delivto i t by the engine so long as an equilibrium

LER TORQUE ABSORBED (AIR RESISTANCE) a1 aircraft installation the Prony replaced by a propeller which absorbs es the torque delivered to the propeller the engine. The distance from the cenessure of the blades to the center of the

CENTER OFAfR PRESSURE

and collar remains a rotating shaft exerts Ib-ft a t 2000 rpm, i t will e same torque a t 2400 rpm. With a propeller, on the other hand, the load is provided by the resistance of the air, and in the case of a fixed pitch propeller the torque absorbed is a definite function of speed. It is found by experiment that, within a wide range of tip speeds, the torque varies directly a s the square of the propeller's rpm; in other words: Torque = K x rpm2 where K is a constant whose value depends on such factors as the density of the air and the number, size, shape, and angle of the blades. In simpler terms : if the speed of a fixed pitch propeller is doubled, i t imposes four times the initial torque on the shaft; or, if a propeller absorbs a torque of 1000 lb-ft a t 2000 rpm, i t will absorb a torque of 1440 a t 2400 rpm, provided the pitch of the blades remains unchanged and tip speeds do not approach the speed of sound.

PROPELLER LOAD CURVE The so-calIed "prop load curve" is a graph which shows the relation between the rpm of a fixed pitch propeller and the power required to drive i t a t a given speed.

g the friction adjustSome engines incorpo-

The equation for the curve is readily derived as follows from the relationships discussed above in the paragraphs on Horsepower and Propeller Torque, thus : Since: Hp = k x Torque X rpm and: Torque = k' x rpm2 by substitution: Hp = K x rpm3 or, expressed as a proportion: Hp, - ( R P ~ , ) ~ HP,

-

(Rpm2Is

In other words, the power absorbed pitch propeller is proportional to the c rpm-a relation which holds true unt speed of the propeller blades approache sound. In less technical jargon, power must be made eight t i double the speed; and twenty-se great to triple it. For example:

.

-

POWER AND ITS MEASUREMENT propellers, the power absorbed is independent of the rpm, for by varying the pitch of the blades the air resistance, and hence the torque, or load, can be changed without reference t o propeller speed. Consequently, with variable pitch propellers a given torque may be absorbed a t an inAnite number of rpm, and, since power is a function of both torque and rpm, a particular horsepower may be developed a t an unlimited number of torque-rpm combinations.

-

ENGINE TORQUE A FUNCTION OF MANIFOLD PRESSURE

PERCENT OF NORMAL RATED R P M

Fig. 20

- Propeller Load Curve

Propeller RPM

HP Required

1000 1500 2000

80 270 640 1260 2160

-

2500

3000

The figures apply, of course, to only one particular type of propeller; larger propellers will require more horsepower for any given rpm; smaller propellers will require less ;but the relation between rprn and hp will in all cases be given by the formula : - - - "... - - - Hp = K x rpms It will be seen from the figures and from the graph that, a s the rprn become higher, the additional horsepower required to produce a given increment in speed (say 500 rpm) increase very rapidly. Only 190 additional hp are required to raise the rprn of the propeller in the example above from 1000 to 1500 rpm; but 910 additional ;hp are needed to raise it from 2500 t o 3000 rpm. ' Since speed is the only means of regulating the 'torque, and hence the power, absorbed by a fixed &itch propeller, it follows that, if the power outc: jput of the engine is changed, the engine will :accelerate or decelerate until an rprn is reached T a t which the power delivered is equal t o the power absorbed. In the case of variable pitch

-

'

OFF

It remains to discover the not very mysterious process by which the torque absorbed by the propeller is produced and delivered by the engine. The process is the familiar one of applying a force a t right angles to an arm, the complications in this instance being that the amount of the force and the length of the arm are continually changing. The force is not a weight or a constant resistance, but a varying pressure-the pressure exerted on the piston head by the gases expanding in the cylinder. This force reaches its maximum shortly after the beginning of the power stroke, and not long afterward acquires, in effect, a negative value during the compreasion stroke. The length of the arm a t its maximum is roughly the length of the crank throw multiplied by the mechanical advantage (2:1, 3 2 , etc.) of the reduction gearing. It is zero whenever the piston is a t top or bottom dead center. However, in spite of the fact that an engine actually delivers its torque unevenly, the frequency and duration of the power impulses, together with the momentum of the moving parts, make it possible in practice to consider torque and with it rprn and power as being continually produced.

fi -I ,

a

'

.

,;

Any factor which affects the mean or average pressure acting on the piston will change the torque. If the mixture strength, temperature, and type of fuel remain unchanged, the pressure developed in the cylinder will depend on the masB ' m of the charge that can be forced into it. This, in turn, will depend on the pressure existing a t the : intake port, which is substantially the same as that a t the collector rim, or manifold pressure. Torque thus turns out to be a function of manifold pressure, and varies directly with i t throughout almost the entire operating range of the engine.

I.

'

PWA. 01. 100

PRA!I"T & WHITNEY AIRCRAFT

TORQUE HERE

WHICH V A R ~ E SWITH MANIFOLD PRFSSUREJ A M /S REGULATED BY THE THROTTLE CONTROL

ENGINE CONTROLS The use of the throttle lever a d the propdler governor control is implicit in the formula : Hp = K x Torque x rpm Insofar a s torque varies approximately with manifold pressure, i t would be possible for purposes of this discussion t o write : Bhp = K x Manifold Pressure x rpm + C where K and C are constants whose values depend on engine characteristics and operating conditions.

1

I I

I 11 1

1

balances the increased torque exerted by the engine-there being no means of regdating propeller torque other than by rpm. Accordingly, with a fixed pitch propeller operating a t a given altitude, it fo-llo~vsthat for each throttle setting, o r manifold pressure, there is a corresponding rpm and power.

Horsepower, or bhp, is thus a function of both manifold pressure and rpm. Manifold pressure is regulated primarily by t h e throttle, supplemented in certain types of operation by auxiliary supercharger controls. Opening t h e throttle on an engine equipped with a variable pitch-constant speed propeller will not change the r p m ; but i t will increase manifold pressure, torque, and horsepower. Power may be increased in this manner without a change in rpm until the airpumping capacity of t h e induction system is reached a t full throttle, or until cylinder pressures become excessive or detonation beginc.

Rpm is controlled by regulating the torque absorbed by the propeller-in other words, by increasing or deereasing the resistance offered by the air to the pmpeller. With variable pitch propellers this is accomplished by means of t h e propeller governor control lever, which adjusts the blade angle. Power may be increased b: raising the rprn, a t constant manifold pressure, until an engine speed is reached a t which reciprocating and centrifugal loads become excessive. In theory the engine should develop i t s greatest power when maximum permissible manifold pressure (torque) is combined with maximum permissible rprn. (In practice, supercharging characteristics and volumetric efficiencies* vary with engine speed, and a change in engine speed is usuaIIy accompanied by a secondary change in manifold pressure if the throt-

In the case of an engine equipped with a fixed pitch propeller, a n increase in manifold pressure is necessarily accompanied by an increase in rpm, until t h e torque absorbed by t h e propeller

"Volu~netricEfficiency: volume of air at atmospheric pressure inducted into the engine every t ~ revolutions ~ o of the engine divided by the latter's total piston displacement.

POWER, BMEP AND RATINGS

BMEP

tle position is left unchanged. For the same reasons maximum power a t constant manifold pressure is not necessarily obtained a t the highest speeds. Furthermore, it must be remembered that because of variations in propeller efficiency with changes in pitch and rpm, maximum propeller thrust horsepower does not necessarily coincide with maximum engine bhp.)

Where fixed pitch propellers are used, the operator has no control over the rpm except, as explained, by means of the throttle (i.e. manifold pressure). With the qualifications mentioned parenthetically above, the engine will develop its greatest power when manifold pressure is a t the highest possible figure, since this will correspond with the maximum rpm.

BRAKE MEAN EFFECTIVE PRESSURE (BMEP) One of the basic limitations placed on engine operation is imposed by the pressures developed in the cylinders during combustion. On the one hand, a s these pressures become greater, the result is an increase in power; on the other, if they become too great they impose dangerous loads and lead to excessive temperatures which may result in engine failure. It is important, then, that the operator have a t his disposal some means of estimating these pressures, not only t o enable him to protect his powerplant, but also to direct him in the efficient application of its power. Cylinder pressures are not recorded on any cockpit instrumen-t or gage, and their direct measurement is actually very difficult. They are related to manifold pressure, since any increase in the latter results in a heavier weight of the charge being forced into the cylinders, and, consequently, a greater amount of energy being liberated by combustion. However, a number of other factors must be taken into account, notably engine speed, and, a t best, manifold pressures are merely qualitative measurements. The most convenient way of estimating cylinder pressures

proves to be an indirect one, and the index of measurement most commonly used is the brake mean effective pressure, or bmep - a term that isglibly bandied about more frequently, perhaps, than it is understood.

CYLINDER PRESSURES A series of complicated and elaborate measurements indicate that the pressures actually existing in a cylinder of an aircraft engine during a complete power cycle can be represented by a graph similar to Fig 21.

INDICATED HORSEPOWER (IHP) The shaded areas, P and I, below the curve represent the power developed in the cylinder: P, by the combustion of the fuel-air mixture; I, by the pressure of the incoming charge. The unshaded areas, C and E, represent the power expended: C, in compressing the charge ; E, in expelling the exhaust. The net useful power, or indicated horsepower (ihp) , developed may thus be represented by the difference between the shaded and unshaded areas, or:

+ +

Ihp = Shaded Areas (P I ) - Unshaded Areas (C E)

PfS fCEN rRAYEL

Fig. 21 -Cylinder Pressures During Power Cycle

PRATT Br WHITNEY AIRCRAFT

PWA. 01. 100

INDICATED MEAN EFFECTIVE PRESSURE (IMEP) AND ACTUAL PRESSURES

The relationship between imep and the actual shape of the pressure curve has been established by experiment, and the range of incipient detonation thoroughly explored. Accordingly, if t h e imep of an engine can be measured, significant peak pressures and other combustion characteristics revealed by the curve can be predicted with sufficient accuracy for practical purposes.

%

I t is sometimes convenient to think of t h various cylinder pressures-useful and adverse -21s combined into a single uniform pressure which acts during the power stroke alone, and is capable of producing the given indicated horsepower (ihp) . This may be illustrated graphically by drawing a constant pressure line, AB, for the power stroke alone a t such a distance above the horizontal axis t h a t the area enclosed by the rectangle, AECD, is equal to the net area representing the ihp, thus: Area ADCD = Shaded -4reas (I' + 1) Areas (C + E ) = Ihp

-

UA1,ClTLATION OF IRTEP

If the indicated horsepower of an engine is kno\vn, its imep can be computed from the formula :

Unshade' Tmcp (psi)

The mean, or average, pressure (AD or BC effective during the power stroke is known a s the indicated mean effective pressure (imep). The operator is naturally more concerned with actual shape ancl peak pressure of the curve than with the abstract notion of indicated mean effective pressure (imep). For instance, a curve which rises steeply t o a n exaggeratedly high peak and descends thence in rapid oscillations, while it may enclose substantially t h e same "power area" a s a smoother, flatter curve, indicates that the power of combustion is being delivered in such a way that it cannot be readily absorbed by the engine, and may prove harmful to the latter. The detonation curve below is an example,

=

792000 Displacement

ihp rpm

ihp - rpm

= K --

he derivation of this formula is briefly a s follon-s: The force, ~n Ibs, exerted in any cylinder is the product of the imep (in psi) by the area (in sq in) of the piston head on which it acts, or: F ( l h s ) = imep (psi) y A ( s q in.) The distanre, in feet, through which this force mores during the power stroke is the length of the stroke (in in.) divided by 12. or: ~ ( f t= ) S(in.) I12 The work, in ft-lbs, done cluriqg the power stroke i s Work (ft-lbs) = F

x TI = i ~ n e px

A x S112

The displacement, in cu. in., of an engine i s the product of the area of the piston ( A ) by the stroke ( S j by t h e number of cylinders ( N ) . Accordingly: Work per rev. =

iniep

x Displacement 24

The work per minute is the work per revolution times the rpm, o r : Work per rnin =

imep

x Displ. x rpm 24

D i ~ i d i n gthe v o r k (still In ft-lbs) per n ~ i nby 33000, we obtain the ihp, so: Ihp = fC NORMAL COMBUSTION

b

Fig. 22

i n ~ e px Displ. 24

B. C OZTONATION

- Cylinder Pressures

Normal and Detonating

x

x rpm

3 3 0 0 0-

Transposing the knc~wn quantities to one side of the equation we obtain: Imep

792000

ihp

Dlspl.

rpm

= TX-----

BMEP

POWER, BMEP AND RATINGS And, unless an engine sloughs off a cylinder, its displacement remains constant, and the formula reduces to:

(Thus, for a Twin Wasp D engine: K =

792000 2000

= 396)

This would be a satisfactory way of computing imep, and thus predicting actual cylinder pressures, were it not for the fact t h a t ihp is difficult t o measure directly.

CALCULATION OF IHP (IHP = BHP + FHP) Not all the power developed in the cylinders (i.e., t h e ihp) is available a t the propeller shaft.

As suggested in Fig. 23, some of i t is lost in overcoming internal friction, some is diverted t o driving the accessories, such a s the magnetos and fuel pumps, and some is absorbed in driving the supercharging mechanism. The total power thus lost is known a s friction horsepower (fhp). What remains in the form of useful power delivered to t h e propeller shaft is called brake horsepower (bhp). Ihp = bhp

This relationship makes possible a sufficiently accurate and relatively simple computation of ihp, and suggests a short-cut to t h e determination of cylinder pressures. Bhp can be readily calculated by means of a Prony Brake, a dynamometer, or torquemeter. In fact, whenever the relation between bhp, rpm, and manifold pressure has been calibrated for a given engine, the bhp may be found from a power curve, if the last two quantities are known and t h e proper corrections made for carburetor air temperature (c.a.t.) . Friction horsepower (fhp) is not a constant quantity, but increases with the rpm and impeller speed or stage of supercharging. It can be determined for any given rpm and supercharging combination by measuring, on a modified dynamometer, the power necessary to turn over the engine with t h e cylinders not firing. Such a calibration is normally made for each type of engine, and covers its entire speed range. With bhp and fhp thus known, ihp is readily calculatecl, and from i t imep and cylinder pressures may be determined. BMEP-Definition

and Computation

In practice, however, actuaI cylinder pressures are not determined from ihp, but from bhp. By analogy, imep can be thought of a s the sum of two pressures ; one going t o produce the power necessary t o overcome friction (fhp) , t h e other t o produce t h e net power delivered a t the propeller shaft (bhp). The first i s called friction mean effective pressure (fmep) ; the second is known by the more familiar, but no less formidable, name of brake mean effective pressure (bmep) .

BHP

m

Fig. 23 - P o w e r s

+ fhp

and Pressures

Since neither bhp nor bmep take any account of friction, the engine can be thought of, in this connection, a s a frictionless power plant in which all the power developed by t h e bmep within the cylinders is delivered without loss

C

PRATT & WHITNEY AIRCRAFT as bhp to the propeHer shaft. Substituting bhp for ihp, bmep can be computed by an equation identical in form and derivation with the one previously used to compute imep, namely:

PWA. 01. 100 than imep, and may represent as little as 8556 of the latter.) Furthermore, the actual peak pressure, as shown on cylinder pressure curves, are in turn greater than the imep. Nevertheless, this abstraction proves to be a useful one in determining the actuaI pressures and in establishing operating limits.

BMEP AN INDEX OF CYLINDER PRESSURES To begin with, bmep may be readily calculated a t any stage of operations in terms of three quantities that are easily found : bhp (from a power curve, torquemeter, Prony Brake, or dynanometer), rpm (from a tachometer), and a constant (792000/DispIacement). Moreover, bmep besrs a definite relationship to imep (bmep/imep = bhp/ihp), which is established throughout the engine's operating range by the calibration of the fhp (ihp I= bhp fhp). Since the actual cylinder pressures and other combustion characteristics can be deduced from imep, and since the relation of bmep to imep is known, it follows that these pressures and other conditions can be determined from bmep as well a s from imep, and the fact that bmep can be more readily computed suggests its use as being a more convenient index by which to rate or limit an engine's performance.

+

Displacement II

rpm

and, if this bmep formula be divided by the corresponding imep formula, it will be seen that: bmep imep

- = -

bhp ihp

Bmep, therefore, proves to be something of an abstraction-a fictitious pressure which has no real existence in the cylinders. Imep is clearly greater than bmep, since i t comprises both bmep and fmep. (Depending largely on the rpm, bmep may be from 10 to 50 psi less

Since a rise in bmep will normally mean an increase in the actual cylinder pressures, structural considerations will dictate that operation be confined below certain bmep values, or the stresses imposed may result in serious failures. Since high cylinder pressures will normally be ompanied by high cylinder head temperares, it follows that bmep values also set a ries of limits for preventing detonation, for rotecting the materials of the cylinders gainst excessive temperatures, and even for electing a type of spark plug. Bmep limits will not necessarily be the same throughout the engine's operating range, for i t is the imep, not the bmep, that actually imposes the mechanical stresses and high tempera tures. If, for example, a two-stage, two-speed engine is run a t high speed in the high auxiliary stage, the fhp will be considerably greater than if the engine is run a t low speed in the main stage (or "neutral"). Accordingly, a smaller portion of the total imep will be avanable as bmep in the first case than in the

POWER, BMEP AND RATINGS second, and the bmep limits must consequently be lower in the former than in the latter, if safe imep, and hence actual peak pressures, are not t o be exceeded.

USE OF BMEP LIMITS IN OPERATION From the relationship : Btnep

bhp =K rpnl

i t can be seen t h a t bmep varies directly a s the bhp and inversely a s the rpm. ble t o run a n engine a t consistent with propelis thereby minimized, iminished, wear reuence of low impeller fuel-air charge dee most efficient retained by operating a t recommended bmep limits, bmep, the greater the bhp m - or, t o put i t animum bhp will be secured m. This type of operation ble t o the cruising range, reases in power are anticiult in exceeding safe As an illustration of t h e foregoing: let us -1980 engine wishes 00 a t the most efficient d economical rpm. Reference t o t h e approprispecifications informs him that for cruising low impeller ratio with 100 octane fuel the ximum recommended bmep is 150 psi. Aping the formuIa: Bnlep = K

x

bhplrpm,

mbinations a s t h e following:

B

Since fha decreases w other t~

-.

two combinations appear t o offer temp sibilities in the matters of fuel econ prolongation of engine life. However, volve operation in high propeller pitch to obtai t h e necessary thrust and torque, and quently impose a heavy load on t h e slo volving engine. This calls for high pressures, a s shown by the corres bmep, with resultant stresses on pisto cylinder parts, and high head temper with t h e possibility of detonation. Cran loads become excessive, and the wear a t thes points may actually be rpm. To avoid flogging turns to the high offered a t the othe run into another set rpm result in grea bhp of 600 is t o be maintained, the ihp be increased, and with i t fuel consum since i t is ihp, not bhp, t h a t ultimately fuel consumption. High rpm also lea cessive reciprocating loads; ant1 the idly a n engine turns over ("other thi equal"), t h e sooner i t will wear itself fly apart. A t the same time high i speeds may also mean excessive char peratures and possi ation clearly lies extremes. Tests made by the manufacture the engine can be safely operated ing range (say, 500-800 bhp), in low impeller ratio with 100 octane fuel, for prolonged periods with the cylinder pressures and texperatures represented by a maximum bmep of 150 psi. Operating above this bmep imposes excessive loads and temperatures on a struggling engine; operating below i t b a s t e s fuel and shortens engine life-in either case t h e consequence may be a forced landing a t some difficult a n uncongenial spot short of destination. The pilot

PWA. OI. 100

PRATT & WHITNEY AIRCRAFT lmum rpnl are very real, not only because fhp, engine wear, and fuel consumption are reduced, but also because low impeller speeds result in low charge temperatures and make i t possible t o "pu:l" high powers from the engine with little chance of detonation. On the other propeller efficiency varies with rprn and angle, and with a particular type of proe r i t may be found more efficient to operate , 1700 rprn with a bmep of approximately i t o prodizce the desired 600 bhp. Other rations may also dictate cruising operalow the maximum recommended bmep ve its col-responding rpm.

, MANIFOLD PRESSURE, AND P CHANGES AT PART THROTTLE an engine is operated a t part throttle and bove the idling range, the volume y the pistons is large; however, partially closed position of t h e valve, the mass of air flowing t o t h e estricted and manifold pressure is umping action of the impeller, under umstances, is secondary to that of the If the pilot reduces engine speed, t h e rate of the pistons is lowered. Alure drop across the valve will t h e throttle setting remains valve will offer relatively less t o t h e reduced airflow called for by e lower rpm. (The mass flow through a fixed orifice varies a s the root of the pressure drop.) Consey, manifold pressure will tend t o rise a s

engine speed decreases - a phenomenon t h a t can be observed, for example, during the ground check of a variable pitch propeller. Inasmuch a s bhp is a function of both rprn and manifold pressure-that is: bhp = K x rprn x Manifold Pressure C-it follows t h a t the decrease in bhp is not proportional t o t h e decrease in rpm, if t h e throttle setting is left unchanged, since manifold pressure is not constant but is actualIy increasing.

+

From the general formula: bmep = K x bhp/rpm, i t will be seen that, if rprn decreases more rapidly than bhp, the value of the fraction bhp,/rpm increases, and a s i t does so the value of t h e bmep also rises. Accordingly, when operating a t part throttle, it is good practice t o observe the following procedure. 1. Reduce manifold pressure by means of the throttle lever before decreasing rpm ;

2. Conversely, increase rpm before raising manifold pressure.

Failure to observe these general rules may lead to excessive cylinder pressures when operating near t h e maximum permissible bmep limits, or to detonation if t h e fuel used is of low anti-knock value. RegardIess of conditions, they constitute sound operating practice. These considerations do not apply t o full throttle operation, since, when no restriction is offered to airflow, bhp varies more or less directly with rpm, with the result t h a t t h e bmep remains substantially constant.

RATINGS NEED FOR OPERATING LIMITS 4 n engine may be made t o last forever. By "pickling" it in a suitabIe preservative, encasing it in a proper container, and storing it in a controlled atmosphere, it should be in a s good condition after several years a s a t t h e beginning storage. Hotveve~,t o serve a useful purpose t engine must be run, and in being r u n it is subjec to wear which will inevitably limit its life The previous discus various forces set up i11 of producing po

MAXIMUM GONTINUOU

e control of ento prevent the setng factor in estaber pressure or rpm been established the tically defined. Engine restricted by limits imoil temperatures, and on tion within which the acty can be obtained are Operation within the ility, as amply demonby the type or proof tests required by curing or regulating agency. Operation side of these limits cannot assure this relility, and the pilot is relyjng on margins of ety that have not been proved. As reliability occupies a place of the greatimportance, and as the ratings define the per limits of reliable performance, they are a ogical beginning for a discussion of engine The ratings of a particular engine are to be ound in the engine specifications and in the ns, and in addiplicable operating ple, to the Specific Double Wasp CAI8 ine, the ratings will check chart as folESGISE SETTING

Impeller Katio Tab-nll

2400

Jfixture Control Poqition

1000

Low

Auto Rich

500

Low

Auto Rich

Low High Low High

Auto Rich Auto Rich Auto Rich Auto Rich

JIax. Jlan. En- Presplne sure Kpm In. Hg 2800

56 5

>larimurn C r u i v Power: Jfaximum Cruise power shall be approved in nrltlnr atrrr con~u1t;ltlonn ~ r hl'ntl & N'hitne Aircraft. T h c urc ol lean nilrrurr (or C o n ~ ~ n u o uCruiw .i shall be r w r v c r r l o r conditions of oprration p r r m ~ t t i n sthe n~ainrrnancr of niaxinlurn hrad temperaturn of 1-5 than JSO Y and 011 inlct ~ r q i p r a l u r wof Ire* Ih,n 185 F. T h e u-c of h i ~ himtiall not IK ~ t t e n l p r e dI [ rhc c ~ r h u r e r o rair Icnipcr~lurec ~ c e e d s

The rating defining the available for continuous operation i Continuous power. This rating is to emergency flight conditions requiring ma sustained output. Automatic rich mixture, equivalent, is required when using Max Continuous rating. NORMAL RATED POWER

This rating is specified by Pratt & W Aircraft and is the maximum recom output for all normal operations such as climb and high speed level flight. Automa rich mixture, or its equivalent, is required w using this rating. In some instances, Normal Rated power Maximum Continuous power are the same. MAXIMIJM CRUISING POWER

Maximum Cruising power is an arbitr established limit for use with automatic mixture, or its equivalent. Maximum Crui limits which apply to all operators have established for the Wasp Jr., Wasp, and Wasp SlC3-G engines. In the case of th models manufactured by P r a t t & Whit craft, the Maximum Cruising power 1 mined individually for each operato obtain the optimum balance between overhaul periods and the desired aircr formance. This limit is establishe'd by tation between the operator and Pratt ney Aircraft after a review of all the affecting the conditions under which gines operate. Individual limit increas follow favorable experience, if desire pourer has been regarded as. 100% erence to this rating. F has meant 50% of No


e policy of establishing cruising ratings ally applies only to the Twin Wasp 'D' r model engines. For previous models, actured by Pratt & Whitney Aircraft, ruise ratings approximated the general f 0.30 bhp per cu. in. for the single row nd 0.375 bhp per cu in. for the double . Ratings, based on these ratios e a satisfactory limit for all normal, re cruising and have given an acbination of durability and perlieu of a positive definition of the ruising rating, the use of limits on the above ratios will insure that conpowers are used pending receipt of formation. In all cases, operase limits will insure the use of are in the range of economical fuel hen the Maximum Cruising rating is specifid, the maximum rpm for continuous lso be indicated. It is not intended ne speed shall be used for all cruisthe most satisfactory results by reducing the rpm as the outLow engine speeds will give rns in the form of fuel economy rability, provided that a proper ance between power and rpm is maintained. he low limit of rpm for any desired power spl-essed in terms of bmep. While the ngs may not specifically mention a bmep tation, experience has indicated the desirity of adhering to a definable relation ben power and rmp. I t is suggested that; the ator base his power and rpm combination rding to the method outlined in Part 111, INGS FOR TIME-LIMITED

s with most mechanical devices, an aircraft ne is capable of sustaining an overload porarily without impairing its reliability without appreciably affecting its duray. Accordingly several ratings have been hed setting up limits within which the be safely operated for restricted a t powers and speeds greater itted for maximum continuous hile take-off demands the output, i t is a condition with

PWA. 01. 100

definite time limits. To obtain this output on the basis of a continuous rating would require a much heavier engine, and the aircraft's efficiency is accordingly increased by establishing a rating which permits the engine to meet take-off requirements, but limits this performance to a short period of time. The specific time limitations are determined by the Military Procurement Agency of by the Civil Aeronautics Authority. As this period limitation is subject to change, the specific operating instructions must be consulted for information regarding a given engine model. For engines with two-speed or two-stage superchargers the Take-off rating is available only in the lowest impeller ratio or stage of supercharging. Exceptions to this rule may be found with individual engine models. For information regarding any one model consult the applicable specific operating instructions. 2. Military Rating. Tactical and combat requirements necessarily dictate the amount of pourer that will be used in military operations. In meeting these demands military personnel are naturally less concerned with the life of the engine than with the offensive or defensive advantage of overload performance. The military rating establishes the maximum limited period performance for service aircraft, and is not available to other operators. In the lowest impeller ratio or stage of supercharging Military Power and rpm are usually the same as for take-off.

An aircraft engine can actually be run continuously under overload conditions of power and speed for much longer periods than those permitted by the ratings. However, the period of reliable operation is thereby reduced to an impractically short time. By imposing a time limit on Take-off and Military power ratings the cumulative effect of the overloads is distributed evenly over the period between overhauls, and the useful life of the engine accordingly lengthened. 3. Maximum Diving Speed. This rati tablishes the maximum safe over-speed dictated by the allowable recipro centrifugal loads resulting from t h which may be encountered maneuvers. Maximum divi limited to 30 seconds.

POWER, BMEP AND RATINGS &-

ADDITIONAL RATINGS I. Temperature and Pressure Ratings. In addition to the power and rpm ratings which govern engine performance, limiting conditions of temperature and pressure are specified by the engine manufacturer, and the observance of these limits is mandatory. These ratings are : a. Cylinder Head Temperature-maximum b. Oil (Inlet) Temperature--maximum and minimum c. Oil Pressure---maximum and minimum d. Fuel Pressure--maximum and minimum Each of these ratings is given for various o p erating conditions. The consequences of failure to observe these limits have already been discussed. 2. Combat Rating+ The conventional Normal Rated, Take-Off and Military power ratings are based on peacetime concepts of engine reliability and durability. In wartime these considerations are not necessarily the first importance, and tactical and combat demands warrant the use of powers in excess of those permitted under normal circumstances. The performance available on such occasions is entirely regulated by the military services, and the demonstration of the engine's suitability for this power output is not an obligation of the manufacturer. Operations of this type come under the category of combat ratings, and are not t o be considered outside of the military services.

RATINGS ators will naturally select powerplants that are permitted t o deliver a high percentage of their maximum potential output. On the other hand, over-enthusiasm regarding ratings is held in check by the consideration that the engine must pass a rigid proof test, and the manufacturer's reputation will suffer if the reliability of his product fails to measure up to the claim made for it. The performance offered is thus the manufacturer's best judgment of the output an engine can deliver, within the accepted standard of reliability and as demonstrated by the type test. ENGINE DURABILITY Durability is the measure of engine life obtained while maintaining the desired reliability. The fact that an engine model has successfully completed its type or proof test is an indication that i t can be operated in a normal manner over a long period before requiring overhaul. However, no definite time interval between overhauls is specified or implied in the rating. The durability realized is determined largely by the type and condition of operation. The period between overhauls may be represented roughly by Fig. 26.

3. 'Other Ratings covering special applications may be supplied from time to time for individ'ual engine models.

RELIABILITY STANDARDS Standards of engine reliability are agreed upon by the engine manufacturer, on the one hand, and the procuring or regulating agency on the other. The accumulated experience of the past makes possible the establishment of testing standards that accurately demonstrate the ability of an engine to perform a t given ratings a t the standard of reliability specified or agreed upon. Competition gives every incentive to the manufacturer to offer the highest possible ratings. Over-conservatism in rating engine performance will result in long engine life-but ften on the manufacturer's shelves, since oper-

TAKE-OFF

0 WWER USED CONTlNUOUSLY

Fig. 26

-Effed of Continuous Power on Durability

The values shown in Fig. 26 are only approximate and vary even a t the same percentage of power output because of such other factors as

BRAKE HORSEPOWER

Fig. 25

- Twin

Wasp Ratings

fuel. The Twin Was

I by the contint eilyines a r e the c a b OF FillEL

me basic Ferent r, " :he grade or" rue1 to be u between these otherwise iden the \.ariations in the carbur

.

lress detonation. For example

:

fuel only. The Twin

o r set sufficietltly 1 ;hese ratings to be developed

ed for the carbure This practice nras for the fighting istirlctly an emer f the fact that engu

~ n c l cnunil~crdoc' not adcquatel\ clc+cr~l~c thc anti-Lnn

ni\turc anti-knocl. n t l n c .\ tucl l ~ k eth.~L Iw de9icnatc:d a- S i 101 Irich rutinc. Thc iollo\r

rsepower, the misture strength p r

r i d or rcrn l ~ l.c. i i c ~ r

PART III

GENERAL OPERATING INSTRUCTIONS STARTING

7 . .

er of methods for obtaining ith any engine model. Most of ry considerably in detail. Some will sults for one individual but fail to action for others; while some methbe used safely only by the most experi-

owed exactly and withThey are based on the hitney Aircraft, and are they can give the test protection to

cedures require

the installation so that the various control nipulations will be accomplished effectively without hesitation or lost motion. On engines using the pressure injecti buretor, Pratt & Whitney Aircraft that the start be accomplished w

recommendation. I t is recom

TT & WHITNEY AIRCRAFT

PWA. 01. 100

general, three different starting procewill cover all Pratt & Whitney Aircraft es; one for engines using float type carbus and the other two for engines using presinjection carburetors. These latter two ocedures vary basically because of differences ming systems. The specific procedures for of these starting procedures follow ; the ious steps are discussed in detail on Pages articular engine, consult the apific Operating Instructions. TING PROCEDURE: A at Carburetors:

-

-Engines

with

Off

Full lean or idle cut-off - Hamilton Standard counterweight type (not Hydromatic) at low rpm (high pitch) ; others at high rpm (low pitch) - Cold - Unfiltered (or off) - Full open - Closed -

retor air heat retor air filter

- '/s open with electric direct cranking, inertia or combination inertia-direct cranking starters. % open with cartridge starters

1. Pull propeller through in the direction of normal rotation, or "inch" engine with starter, for a t least five revolutions of crankshaft:

2. Turn on fuel supply from suitable tank.

3. Move mixture control to full rich. 4. Operate wobble pump slowly, or turn on electric auxiliary fuel pump momentarily, until fuel pressure registers 3 psi.

5. If inertia type starter is used, begin energizing starter while priming. 6. Prime: a. If carburetor has integral primer : (1) Move mixture control to full lean. (2) Pump throttle through complete travel for number of strokes necessary to give required prime.

(3) Return mixture control to full rich.

(4) Operate wobble pump slowly or turn on electric auxiliary fuel pump momentarily until fuel pressure registers 3 psi.

b. If separate primer pump is used: (1) Keep mixture control in full rich. (2) Prime as required.

7. If inertia or cartridge starter is used, turn ignition to "Both On." 8. Maintain 3 psi fuel pressure with wobble pump or turn on the electric auxiliary fuel pump. 9. Engage starter (and ignition booster, if separately controlled).

10. If direct cranking starter is used, wait for engine to turn two revolutions. Watch for indications of "hydraulicking." If none, turn ignition to "Both On." 11. After engine fires adjust throttle to hold engine speed a t 600 rpm. Note: Watch oil pressure gage. If oil pressure does not register within 10 seconds, STOP engine and investigate.

12. If equipped with Hamilton Standard counterweight type (not Hydromatic) propeller, move control to high rpm (low pitch). 13. When oil pressure shows, advance throttle to obtain 1000 rpm. If a start is not made in a reasonable time, i t is possible that: 1. Engine is overloaded (over primed) as indicated by discharge of fuel from carburetor drain, or, in the case of a cold engine, by the presence of liquid fuel in the exhaust outlets of the primed cylinders. In this case :

a. With direct cranking starter: (1) Continue cranking. (2) Place mixture in full lean. (3) Discontinue operating wobble pump or turn off the electric auxiliary fuel pump. (4) Fully open throttle.

(5) After eight revolutions re tle, place mixture control and repeat starting proced

GROUND OPER Note: I t is frequently possible to effect a start while clearing out the engine, in which case the operator must be ready to retard the throttle immediately and move the mixture control to full rich.

b. With inertia or cartridge starter : (1) Turn off the ignition. (2) Fully open throttle. (3) Pull propeller through about 5 revolutions in direction of normal rotation to clear engine. (4) Retard throttle and repeat starting procedure. 2. Engine is under-primed. In this case;

a. Repeat priming procedure to give engine additional prime. b. Repeat starting procedure. possible to start the engine, buretor for malfunctioning. DILTRE : B.-Engines with arhuretors and Using Cylriming.

- Idle cut-off - Hamilton Standard counterweight type (not Hydromatic) at low rpm (high pitch). Others at high rpm -

Single-Stage-low ; Two-Stag-main or neutral Turbo-off

-

Unfiltered (or off) open

- Full

ough in direction of nor" with starter, for a t of crankshaft. suitable tank.

(1) Operate wobble pump to mai 3 psi fuel pressure or turn on tric auxiliary fuel pump. (2) Prime as required. b. If electrically operated priming val used., (1) Operate wobble pump to m required fuel pressure or electric auxiliary fuel pum (2) Prime as required. 5. If inertia or cartridge starter is u ignition to "Both On." 6. Operate wobble pump to maintai fuel pressure or turn on electric au fuel pump. 7. Engage starter or fire cartridge. 8. If direct cranking starter is used, engine to turn two revolutions. W indications of "hydraulicking." If non turn ignition to "Both On." 9. When engine fires, move mixture control

immediately to automatic rich position. 10. Maintain fuel pressure with wobble pu or electric auxiliary fuel pump until e pump builds up specified fuel pressu 11. Adjust throttle to maintain about 600 until oil pressure shows. Note: Watch oil pressure gage. If oil pressure does not register within 10 seconds, STOP engine and investigate.

12. If equipped with Hamilton Standard terweight type (not Hydromatic) p ler, move control to high rpm (low p 13. When oil pressure shows, adjust th to 1000 rpm. If engine does not fire almost immediately 1. With direct cranking starter (except models using updraft carburetors) a. If engine is under-primed : (1) Continue cranking. (2) Maintain fuel pressure.

(4) Return mixture con

PWA. 01. 100


b. If engine is overloaded: (1) .Continue cranking. (2) Place mixture control in idle cut-off. (3) Open throttle wide. ) After about 8 revolutions of the

engine retard throttle and repeat starting procedure, if engine has not started during this clearingout procedure. 2. With inertia or cartridge starter: a. If engine is overloaded: (1) Keep mixture control in idle cut-off. (2) Discontinue operation of' wobble pump or turn off electric auxiliary fuel pump. ) Turn off ignition. ) Open throttle wide.

(5) Clear engine by. pulling propeller through 8 revolutions. ) Retard throttle and repeat starting

procedure. . If engine is under-primed: (1) Turn propeller forward 1/2 revolution to disengage starter jaw.

(3) Repeat starting procedure.

r t is not obtained after a reasonable r of attempts, an investigation should be to determine the cause. in the coldest weather the starting "C" following is recommended as erable for engines having electric and equipped with direct crankation inertia - direct cranking

-

OCEDURE: C. Engines with ection Carburetors and with Blower Throat or Carburetor

ter engines are provided with er throat or carburetor prime priming systems introduce into the blower rim or the oat instead of the cylinder intake priming flow is generally much n with cylinder intake port prim-

ing, and all the cylinders receive a combustible fuel/air mixture rather than just a limited number of primed cylinders. Consequently, the undesirable practice of resorting to the use of the mixture control to supplement the priming fuel flow is eliminated. Instead of priming before engaging the starter, the priming switch is closed simultaneously with the starter engaging switch. As a combustible fue'l/air mixture range is reached the engine will fire, and should be accelerated with the primer, after which the mixture control is slowly moved out of the idle cut-off position to automatic rich, using prime as required until the start is secure. To avoid overloading a warm engine, flicking the primer switch is advisable. A colder engine will require constant priming until i t fires. The elimination of priming before cranking and the use of the mixture control to supplement the prime has eliminated or reduced the occurrence of "hydraulicking" and other hazards associated with over-priming.

A typical starting procedure follows: Control Position Cheek Ignition

- Off

Misture

- Idle

Propeller

- Hamilton Standard counterweight type (not Hydromatic) at low rpm (high pitch) ; others a t high rpm (low pitch).

Supercharger (when applicabIe)

- Single stage -low

Carburetor air heat

- Cold

cut-off

Two stage- main, or neutral Turbo-off.

Carburetor air filter

- Off (unfiltered)

Cowl flaps

- Full open

Oil cooler shutters

- Closed

Starting Procedure 1. Pull propeller through in direction of normal rotation, or "inch" with starter, for a t least five revolutions of crankshaft. 2. Turn on fuel supply from suitable tank. Note: Keep mixture control in idle when the engine is not fir

3. Turn on electric auxili

con t nece

tic rich using prime as a r t is secure. ressure with auxiliary fuel ne pump builds up specified

LO. Ad uni pressure gage. If oil pressure does within 10 seconds, STOP engine and

Hamilton Standard counnot Hydromatic) propelto high rpm (low pitch).

continuously for m shouId be allowed t second start. 3. The Cornhination Starter, as the name i electrically energized ( and a direct cranking

not continue to run : ontrol back to idle cut-off purposes.) As the energy of the fl sipated, the starter motor continu king and priming. ned within a reasonable hould be made to deter-

is supplied in the same manner as fo cranking starter. 4. The Cartridge Starter develops the combustion of a slow-burning tained in a cartridge which is inserted i

1. The 11~ ~ e r tStarter ia CO

-

-

.

1ynrheel which can be

PWA. 01. 100

PRATT & WHITNEY AIRCRAFT CUSSION

.

r .r

- -.--f?y. ,

Control Position Check By checking off the position of each engine and accessory controI the operator makes certain that all powerplant units are in readiness for operation, and satisfies himself that no item is in a position to cause damage.

n2-

.

:.

- -.cc-

4

'\

6

- ?.*

q

#

4 .

.

;

'.

A

A

7-

During the starting procedure ascertain that all is clear of the propeIler before turning on the ignition switch. b. Mixture

The ignition switch should be in the "Off" position a t all times when the engine is not running, except as the starting procedure may require, This is essential for the protection of personnel who may be servicing the aircraft.

Propellel

drive during starting. Engines with turbosuperchargers s have the turbosupercharger contr the " O f f position. e. Carl>uretor Heat and Carburet Vith cor

-.

s rule are the Ham_._---

own with the to retract the

trol should not be until after starting ailling oil pressure

& WHITNEY AIRCRAFT

f. Cowl Flaps

During all ground operation of the engine i t is essential that the cowl flaps (where applicable) be fully open regardless of outside air temperature or the temperature of the cylinder heads. The ignition harness, as well as other powerplant items, require all possible circulation of air to prevent serious damage. Closing of the cowl to accelerate the warm-up will to burning of the insulation on the ignition leads.

PWA. 01. 100 h. Fuel Supply

As a safety factor the fuel supply should be turned off when the engine is shut down, and not turned on until preparation for starting is made. i. Throttle I

il Cooler Shutters

The position of the throttle tant factor in consist

GENERAL OPERATING INSTRUCTIONS

GROUND OPERATION-STARTING

many cubic inches of air a t atmospheric pressure regardless of the throttle position. However, fuel-flow is closely related t o the position of the throttle, but in a manner which differs between t h e two basic types of carburetors. (I) Float Carburetors

In the idling range of operation this type of carburetor furnishes fuel only when a definite pressure differential exists between t h e idle discharge and t h e fuel in t h e bowl. If t h e throttle is too f a r advanced, this differential becomes insufficient t o produce the required flow with t h e results t h a t t h e engine cannot properly fire and, in all probability, back-firing will occur. (2) Pressure Carburetors During t h e starting a d irk3.ing period t h e pm8sURe carbumtw f UP.= nishes fuel entiway as tihe rew1t of the position of the idle w h , which is directly linked t o the :th~cMe. Airflow has no effect the quantity of fuel dis&str+. If .the throttle is too f a r advanad, the fuel discharge will be h o g w k 0the amount of airflow t h a t b being drawn into the cylinders with the result that the engine will be overloaded. In setting the throttle opening before starting, the purpose is to secure the proper fuel-air ratio for combustion and t h e throttle opening is affected by the type of carburetor and starter employed. For example, cartridge starters, which t u r n the engine over a t a higher rpm than other types and consequently produce more airflow through t h e carburetor require a slightly wider throttle opening.

2. Clearing the Engine After the previous shutdown, t h e warm residual oil clinging t o the power section surfaces flows downward toward the lower cylinders. Some of this oil seeps past the piston and piston rings, accumulating in t h e combustion chamber. If sufficient liquid is present, the true compression ratio will

be raised and extremely high pressures will be produced when t h e piston of the cylinder is moved downward on t h e compression stroke. These pressures can be raised to such an extent t h a t damage t o the cylinder, piston, or link rod will result. I n ex-

u

treme instances the piston may actually "bottom" against the liquid. This is known a s "hydraulicking" the engine.

To protect the ehgine against t ity, any excess liquid should be u c ; a r - uuc if the engine has stood idle for two hours or more. To clear out the cylinders, pull the propeller through by hand in the normal direction of engine rotation to obtain a t least five crankshaft revolutioms.

k ,.

,

HINT:

Count the number of propeller blaaes mula pulledtothrough, determine using the the number following of blades forrequired to obtain five crankshaft revolutions : No. Prop. Blades pulled to Obtain 5 5xNo. Blades on Prop. Crankshaft Revs.= Reduc. Gear Ratio (1)

=

5 x No. Blades on Prop. x Reduction Gear Ratio (2) ear Ratio in form: 2:1, 3 2 , etc:.e., 211, 312, etc.) or

'

-

If the propeller cannot be reached to be pulled through by hand, turn the engine

(2) Gear Ratio in form:

.500:1, .667:1,

etc.

Thus: In the case of a three-bladed propeller and a 16:9 (or .5625:1) propeller reduction gear, 9 blades should be pulled through to obtain 5 revolutions of the crankshaft.

Do not pull through in the reverse direction of normal engine rotation. "Backing up" the engine will result in pushing the liquid into the intake pipes, where i t will be ready to return to the cylinders on the next intake stroke.

we^ by "iaeBJnd' w5th Oh'sbrter. This L dlOm by engaging and disengaging the shrber rslo tka* the c m n k h f t turn's-but a few m e e s at a time. L

GENERAL OPERATING WSTRUCTIONS


W/TH E W N E NOT RUNNING AN AUX/LIARY PUMP, EITHER WOBBLE OR ELECTR/C BOOS7; MUSTBEUSED TO RAISE ' FUEL ABOVE TANK LEVEL THE CARBURETOR

FUEL TANK

m

While pulling the propeller through or "inching" with the starter, the operator must be alert for any sign of the piston being forced against unusually high compression. This will be evidenced by a sudden resistance when being pulled through, or by a sudden slowing down when the starter is engaged. This is sufficient indication that an excess quantity of liquid is present in the lower cylinders, and any further attempt to turn the engine over will result in damage. -

If the presence of liquid is suspected, remove a spark plug from each of the bottom cylihders before turning the engine over. This is especially important when the engine is provided with high exhaust tailpipes which do not allow drainage from the lower cylinders.

3. Getting Fuel to the Carburetor

Turning on the fuel selector valve opens a path for the fuel from the tank to the carburetor, but unless the tank is located above the carburetor the fuel cannot reach the carburetor. Therefore, it is necessary to apply pressure to fill the lines to the carburetor, and further, in the case of pressure injection carburetors, to apply sufficient additional pressure to discharge fuel from the carburetor into the engine induction passages when the mixture control is moved out of idle cut-off. Inasmuch as the engine fuel pump is not operating, an auxiliary fuel pump is necessarily used. On many installations this is a hand pump, commonly known as a wobble pump. On other installations, particularly for the larger engines, an electric driven auxiliary fuel pump is used. With float carburetors, the wobble pump should be operated slowly until 3 psi is indicated on the fuel pressure gage, or the electric auxiliary pump should be turned on momentarily to obtain this pressure. By this means the lines and the carburetor bowl will be filled with fuel and the air expelled. Continued pumping or greater pressure may result in fuel overflowing the carburetor bowl and creating a fire hazard.

1

PWA. 01. 100

P R A l T & WHITNEY AIRCRAFT

I I

In the case of pressure injection carburetors, the minimum pressure required will engine mode11, and .ed and main1tained without danger, provided the mixture control is kept in idle cut-off until the engine fires.

tion of the prime and engaging the starter. This is desirable in order to minimize the possible effect of raw fuel lying inside the engine. 5. Priming

If the mixture control is moved out of idle cut-off, fuel will be discharged into the induction system. As the engine is not turning over, only the air contained in the supercharger will absorb vaporized fuel. The drain valve will carry away only a portion of the liquid that the carburetor discharges. The balance of the liquid fuel pours into

As the carburetor cannot supply the cylinders with a combustible mixture of air and fuel without sufficient airflow, the initial firing charge must be prepared by other means. The air contained in the cylinders and induction passages and that introduced during cranking must be used in providing the initial firing charge. By spraying fuel into this air, a fuel-air mixture that is within the combustible range is prepared to give the initial firing impulses which turn the engine over a t a speed that will bring airflow and normal carburetion. The fuel is furnished through the priming

the lower intake pipes, paving the way for "hydraulicking." 4. Energizing Inertia Starter

Energizing the inertia starter (where used in conjunction with cylinder intake port priming) while priming the engine results in the least time delay between the comple-

-



TEMPE

F.*. system.

h

27 -I?Viect of Temperature ma

.- .

.

It i's a b m i d by nozzles to assist

tb -\rapcw&&tLonaf the fuel and is discharged hb @bWa&h system or the intake ports @f the bp cylinders of the engine d ing m pn the type of priming system m the ermgirr~ The prime k b l d contain enough eosnplebe4y valp~~&& f d i% fin. %heinduc. .

The ammnt of prim squired vm-ies with Re.

k one &reme af tmm'zam

id frwl mu& be injeetd to &&in this iixd vapor requirmmt. This ig kcawe k m m t u m affects tE.le y a g m r i ~ tiaa af &e fuel. The .-dm$ amat af p r b &. estirn~otd hjt hoking at ttR W0whg g

(1)

Air Tmwratum : TeZYs the hamwmhm of the air that wril,l be

PRA'IT & WHITNEY AIRCRAFT drawn into the engine during cranking and after starting. I t aids in estimating the evaporating capabilities of the cold air stream. (2) Carburetor Air Temperature : Tells the temperature of the air in the duct, as well a s some idea of the carburetor temperature and the contained fuel, and hence gives an idea of evaporating conditions. (3) Oil Temperature: Gives an indication of the stiffness of the engine: whether it has been in operation recently or not, if it has been warm or cold during storage. It also gives an idea of the temperature of the supercharger section and its ability to vaporize the first trickle of fuel. (4) Cylinder Head Temperature : Indicates the amount of heat available

Under extremely priming before cr priming constitutes a serious "hyd

PWA. 01. 100

-

ing the cylinder walls and of piston seizure. If the engine has been overprimed it is essential that fresh oil be sprayed on the cylinder walls before etarting. Dry cylinders may be indicated by a squeaking heard while the engine is being pulled through by hand.

reme cases the excess fuel will p the intake pipes of the primed

dilution and a well developed technique are needed for starts under 40 F (5C). Even a t this latter temperature, preheating of the engine is advantageous, if preheat is available, in order to reduce the amount of liquid fuel remaining after priming. (Priming systems are being developed for some engines which permit consistent starts a t (-20 C ) or less, consul Weather Operation.'? a. Overpriming

With cylinder intake port priming, one serious result of overpriming, or of continued underpriming with unsuccessful attempts to start, is the presence of liquid fuel in the cylinders. This washes off the film of oil orithe cylinder walls, pistons and piston rings. Without this lubrication there is a possibility of scor-

This condition usually results in weak firing which does not have sufficient energy to turn the engine over or else causes backfiring. If individual outlet exhaust stacks are used, there will be no evidence of fuel vapor in the exhaust outlets of the primed cylinders. In cold weather, fuel discharged from the fuel drain does not necessarily mean that the engine has been overprimed as only a small portion of the fuel will be vaporized.

GENERAL OPERATING INSTRUCTIONS C.

Sources of Priming Fuel

Fuel for the priming system is commonly provided by : (1) Priming pump integral with carburetor (float type carburetor) (2) Hand pldnger priming pump (3) Electric solenoid priming valve which is inte-

of the throttle is dependent on the amount of prime desired. (e) Maintain fuel pressure 3 refill the carburetor bowl wi fuel after two or three s t of the throttle. ( f ) Return mixture control t rich. (2) The hand priming pump is conjunction with engine i tions equipped with either fl pressure injection carbureors. --

( a ) Place mixture control in full rich. This allows the carburetor bowl to be vented. (b) Raise fuel pressure to 3 psi. This fills the carburetor bowl se the carburetor to

priming pump is a plunger pump, and is usually located cockpit or possibly in the nacelle. T h e quantity of prim gaged by the number of s t made while priming. To prime: (a) Mixture control in idle cut-off.

A n ' & WWITNEY AIRCRAFT (b) Raise 3 psi fuel pressure with wobble pump or turn on electric auxiliary fuel pump. Pressure is necessary to fill primer pump cylinder on the intake stroke.

PWA. 01. 100 (3) The electric priming valve can be used with pressure injection carburetors, or with float carburetor installations if sufficient fuel pressure is available. The electric priming valve is usually attached to the

(c) While maintaining fuel pressure, draw plunger slowly out to ensure that the pump cylinder fills completely. Force plunger in rapidly in order to atomize the fuel effectively a t the discharge nozzles. Prime t h e r e q u i r e d n u m b e r of strokes.

carburetor or some other source of fuel under pressure in the engine compartment. When the solenoid valve is opened, fuel under pressure is distributed through the priming system to the priming nozzles. The amount of prime is gaged by the pressure a t the priming valve and the length of time it is held open. To prime: ( a ) Keep mixture control in idle cut-off. (b) Rbise the'required fuel pressure with the wobble pump or the electric auxiliary fuel pump.

i


GROUND OPERATI

(c) Close primer switch for required time which is dependent on temperature and the particular method of introducing the priming fuel into the engine. e is very difficult eviously pointed out, ctor. With the electric discharged by the primself-priming carburetor manner of operating them amount of prime delivered. ividual technique of the opthe amount of priming devidual may go through the ting with very exact complete operaay not perform the degree of suret feasible to draw e of priming requirements.

two revolutions, the engine sho to come against a definite abrupt sto may be concluded that the lower cylin have accumulated liquid, and no attempt should be made to turn the over until this liquid has been

7. MIXTURE CONTROL C _ C -

*

,--v

or one to two strokes may be ngine, up to three strokes for a cold engine. Installations e. However, this is dependent on the rburetor integral lectric priming valve discharge wilI vary e fuel pressure and the system of introuel into the engine. Consequently time will vary according to the gine being started.

+he case of inertia or cartridge starters, the ignition must be turned on prior to stai.ter engagement, as the full starting -...~~~-uceclure must be completed while the starter energy is available. The direct cranking starter makes i t possible to delay turning on the ignition until after the engint has be~ en turned over two revolutions. Thi s offers tain ~rlereis no e to CalirP "h yclraulic 111

L L -I.

I

/

L

LL

In the case of float carburetors, the mixture control is already in the automatic rich or full rich position when the starter is en-


ment out of idle cut-off when starting often termines the success of the start. If the ~ x t u r econtrol is moved into automatic rich before the engine has "caught," raw is dumped into the engine before norairflow is set up and the engine bees loaded. As a general rule, firing of engine should be of sufficient intensity roduce a t least 350 or 400 rpm before mixture control is moved to automatic . Moving the mixture control into autoic rich a t the first sign of feeble firing r popping will usually be the cause of an nsuccessful start. However, with some riming systems, if movement of the mixre control is delayed too long after the gine is firing normally, the engine will ave used up its prime and fail to start, and ackfiring will often result. If the priming as been properly made there will be suffint and proper strength fuel-air mixture the engine to keep it turning over a t cient speed until normal carburetion refrom proper m e m e n t of the mixture

engine "dies" or sho~vs gine is stopped a t once. Failhas probably caused more ngines than failure to clear

01 into idle cut-off, thus stopping the

nce to pick up, after which the mixture

on as the engine fires, the rpm should Id dourn to about 600 rpm until the essure registers on the gage and it is certain that the engine is receiving oil. In many instances the oil pressure will not register immediately upon starting. In most installations 10 seconds without oil pres-

PWA. 01. 100

sure indication should not be exceeded. However, on some installations having long pressure lines, or pressure indicating transm{t$ing systems, there may be a time lag requiring that this time limit be lengthened. I t is imperative, however, that strict attention be given the oil pressure gage a t this stage of the starting procedure, and the acquiring of this habit may some day pay dividends in saving an engine. 9. PROPELLER CONTROL

In the case of the Hamilton Standard counterweight type propellers, the control should be moved to the high rpm (low pitch) position after oil pressure is indicated.


for a brief interval as outlined in t h starting procedure. (See page 6 is the only instance where i t is sible to move the mixture contro idle cut-off prior to actual engine tion, and the operator must us to ensure that the time limits indi are not exceeded. Introduction of sive fuel into the engine may cause draulicking" or overloading of the e gine.

OF ENGINE

Overloading of the engine with raw fuel is usually accompanied by a heavy discharge of fuel from the supercharger drain in warm weather or the presence of liquid fuel in the exhausts of the primed cylinders. Weak firing, followed by dense black smoke and possibly fire from the exhaust outlets, is also indicative of overloading.

11. "SAVING" THE START

.

.

111 1~nstallationsincorporating

GROUND OPE

cylinder in-

b. If the start is not accomplished b of overloading, the fuel-air rati be leaned out by drawing more a1 the engine. This is done by opeain throttle wide and keeping the control in idle cut-off while continues. As the fuel-air ratio and passes through the co range, a start should be case, immediate action on the part of operator is required to retard the thro so as not to overspeed the engine, and move the mixture control to auto rich to establish normal carbur It will be seen that the manipulat the mixture control in the event derpriming and of the throttle event of overloading is for the of passing the mixture charge f extreme of the fuel-air other so that a combustible ran be reached and a start effected. In the case of inertia or cartrid ers, failure to prime properly corrected by clearing out the engi repeating the complete starting dure. Whether the Tailure t been due to underpriming or to priming, it is recommended that t gine be cleared out so that the op will know the amount of prime tha the engine at any one time.

take:port priming and direct cranking start-

ers, an opportunity is provided to "save" &I., L U ~ : start with

a. If the engine ditional fuel n

cessful starts. The percentage of succes

GROUND OPERATION

between the engine of an autoof an aircraft has been disforming habits of aircraft enIt is not by any means unsafe t a cold automobile engine, pull away ram the curb and drive away. A few blocks f slow, restrained running will suffice to warm he engine to the point where i t will "take" the ccelerator and function normally. As long a s lo attempt is made during the early stages to ross tracks in froat of trains or to perform 'ancy traffic maneuvers in the wrong lane, the -ar and driver wilI be looked upon with favor by the insurance company. The automobile does not depend upon its engine for sustenta-tion. At the instant the airplane leaves the ground n d until i t is in full flight and clear of

formance and tb make such checks as will give assurance of satisfactory functioning. A cold engine is not a dependable engine.

WARM-UP After starting, several minutes will be spent in warming up the engine. This is done a t 1000 rpm* with the propeller in the high rpm (low pitch) position. One thousand rpm is specified as this engine speed will ensure freedom from spark plug fouling. The propeller pitch position results in the lightest possible load a t this rpm. If power were applied to a cold engine there would be an unsatisfactory response for three reasons : a. Oil:-Cold,

undiluted oil is too thick to

be starved for lubricati a high pressure show sure gage. The oil


GROUND OPERATION-WARM

UP

the design, are properly adjusted. The entire mass of metal must be in the working temperature range to have a harmonious relation between the individual parts.

ALUMINUM ALLOY U S E

SEEL

NCK

m

\,

BEAHNG

SLKR LEAD

than l O O F (40C) before increasing engine speed above the warm-up rpm. b. Induction System :-Engine performance and operation are sensitive t o the temperature of the induction passages.

CHA RCE OF /NG OUT MIXTURE

An intake charge cannot hold all its fuel when i t comes in contact with cold metal surfaces, and this leaning-out will cause the engine t o hesitate and run ragged1y. c. Even Expansion of Entire Engine: The engine designer has established t h e clearances between the working parts after considering the effects of expansion when the materia1 has warmed up. In many instances the cold clearances are not sufficient to allow satisfactory oil flow. Because many major parts are made from dissimiIar metals, they must be brought up to operating temperature in order that the uneven rates of expansion, t h a t have been taken into account in

In general, oil temperature is t h e most direct gage of the temperature condition of the entire engine. If time has been taken to bring the oil temperature up t o 1 O O F (40C) a t 1000 rpm with the propeller in high rpm position, the minimum requirements of a proper warm-up will be met. Where practical, an oil inlet temperature of 140-165F (60-75C) should be reached. 1. Cowl Flaps Do not close the cowl flaps t o accelerate the warm-up, regardless of outside air temperature or the indication of the cylinder head gage. The heated air coliecting around the exhaust system will be confined in the area adjacent to the ignition leads. For the protection of the ignition system, it is essential t h a t the cowl

flaps remain wide open a t all times that the engine is running on t h e ground, and for a cooling-off period after the engine is stopped.

PWA. 01. 100

PRAlT & WHITNEY AIRCRAFT

2. Mixture

type carburetors or with fuel nozzles located a t some distances upstream from the impeller, it is desirable to raise the carburetor air temperature during warm-up to prevent ice formation, to assist distribution, and to ensure smooth operation. Do not exceed lOOF (40C).

4. Carburetor Air Filter (where applicable)

Do not u8e a lean mixturn wcl4bdfaf.i~ warmup. Getually, at the warm-up qm there k ~ a c t k a l l yno d i f l e r m in the mixture sugp1ied to the engin@whether the mixture is in a lean or rich pwitian, a=@ metering in this power range ia gwerad by the throttle pasitia As there is ns accdcm while C& &4 ground to uae e\ lean miarbre pSfti6r1, R ia imperative that the &We WQJ. remain in the rich pcdtsim 4m r n ~ b @re that i t will rra4; be Beak@ h z ~ d t e n prior t o take-ob. Engine exhaarst smoking, whlk u with c h d tlrpott1i3, srhoald be idid'Irwted by proper Sdle adjustment, venting fouled p l u m

Where dust conditions are present the carb.uretor air filter should be used. As soon as the engine has started, the air should be drawn in from this source until after the take-off has been performed and an altitude reached where dust-free air is present.

5. Magneto Safety Cheek

Dwhg the warm-up running, the magneto ~ a f e t ycheck

Carburetor heat can be used a s required under conditions leading to ice formation. In the case of engines equipped with float

can be performed. Its pur-

pose is to ensure that all ignition connections are secure and that the ignition sysbrn d l 1 permit opgmItion at the higher powar used in the sound e h k to be eon-

7

GENERAL OPERATING INSTRUCTIONS ducted later. The magneto safety check is conducted a s follows : (1) R P ~ (2) Propeller

- 1000 - High rpm (low pitch)

All other controls the same as during warm-up (3) Switch

(4) Switch

(5) Switch

-From "Both" to "Right" and return to "Both" -From "Both" to "Left" and return to "Both" -From "Both" to "Off" momentarily and return to "Both"

While switching from "Both" to a single magneto position, for example from "Both" to "Right," a slight but noticeable drop in rpm should occur. This indicates that the opposite magneto has been properly grounded out ,and that the connection to the single operating magneto is secure. Complete cutting out of the engine when switching from "Both" to "Off" indicates that both magnetos are properly grounded. Failure to obtain any drop while in the single magneto position, or failure of the engine to cut out while switching to "Off," indicates that one or both ground connections are not secured. The time required for proper warming-up gives ample opportunity to perform this simple check which may disclose a condition which wouId make i t inadvisable to continue operation until after corrections have been made. 6. Ground Cheek

en given sufficient make certain that work together, i t

GROUND OPERATION-GROUND

CHECK

(d) Is the fuel system delivering fuel to the engine a t the required pressure? (e) Are the propeller and supercharger shift mechanisms and other accessories functioning properly? The answers to these questions are learned in the course of the ground check; A standard ground check is performed as follows: The aircraft should be headed into the wind, if possible, to take advantage of this cooling airflow. 7. Control Position Check Cowl Flaps Mixture Propeller Carburetor Heat Carburetor Air Filter Supercharger Control Position (where applicable)

- Open

- Rich - High

rpm

- Low,

Neutral, or Off

- Cold - As required

Procedure 1. Check propeller according to propelIer manufacturer's instruction. 2. Open throttle to manifold pressure equal to field barometric pressure 3. Switch from "BOTH" to "RIGHT" and return to "BOTH." Normal drop - 50-75 rpm* Maximum drop - 100 rpm* 4. Switch from "BOTH" to "LEFT" and return to "BOTH" Normal drop - 50-75 rpm* Maximum drop - 100 rpm* Maximum difference between "RIGHT" and "LEFT" - 40 rpm* 5. Check: Fuel pressure - 17 psi* Oil pressure - 85 psi* 6 . Note rpm 7. Retard throttIe NOTE: Additional features, such a s multiple speed or multiple stage superchargers, require ground checking. The resulting changes in the above basic procedure will be explained in the supplements describing the operation of these special items.

In addition to the operations outIined above, the functioning of various items of airpIane equipment will be checked in an appropriate order. *These quantities are for illustration, Consult the specific operation instructions for values applying to a particular engine.

PWA. 01. 100

PRATT & WHITNEY AIRCR

engine is capable of delivering a given power a t a given rprn and manifold pressure. The original calibration, or measurement of power, is made by means of a dynanometer. During the ground check, measurement of power is made with the propelIer. With constant conditions of air density, the propeller, a t any fixed pitch position, will always require the same rprn to absorb the same horsepower from the engine. This characteristic is used in determining the condition of the engine. Any increase in manifold pressure required to obtain the same rpm, or conversely, the inability to obtain the check rprn with the check manifold pressure, is an indication that the engine is not giving the performance of which it is capable. Variation in altitude of the fields on which the check is made will result in varying manifold pressures for a given rprn a s is indicated in Fig. 28.

DISCUSSION 1. Propeller Check

The propeller check is performed to ensure proper operation of the pitch control and the pitch change mechanism. Each type of propeller requires a different procedure, and the applicable manufacturer's instructions should be followed. 2. Rpm and Manifold Pressure

Before starting the engine, observe the manifold pressure gage. This gage will read approximately the atmospheric (barometric) pressure when the engine is not running. At sea level this is approximately 30 in. H g and a t fields above sea ievel the atmospheric pressure will be less, depending on height above sea level.

Specific check of rprn and manifold pressure relationship should be made during each ground check. This may be done a t the time the engine is run-up to make the magneto check. The basic idea of this check is to measure the performance of the engine against an established standard. Calibration tests have determined that the

When the engine is started and is accelerated the manifold pressure will fzll off until about 1600 or 1700 rprn is reached when it wilI begin to rise. At approximately 2000 rpm, with the propeller in

a

'OLD PRE. YE NEED 1

FAILURE TO 087;4lN CHECI EQUAL TO ATMOS'PHERIC P

z I

U

a

3

I I

STABLISH BY YGINE COMa t r u n ANU LOW PITCH BLADE ANGLE EMPLOYED.

I

200

400

800

800

1000

1200

R P M -PROPELLER

1400

1600

1800

2000

2200

I N FIXED LOW PITCH

Fig, 28 -Manifold Pressure vs Rpm

- Propeller in Fixed Low Pitch

2400

2600

2800

GENERAL OPERATING INSTRUCTIONS fixed full low pitch position, the manifold pressure should be the same a s the atmospheric pressure. That is, if the manifold pressure gage reading (atmospheric pressure) is 30 in Hg before starting the engine, the pressure reading should return to 30 in. Hg a t approximately 2000 rpm, or, if the manifold pressure gage reads 26 in. Hg before starting, i t should read 26 in. Hg again a t approximately 2000 rpm. The exact rprn may vary with various models of engines or because of varying propeller characteristics. In certain installations the rpm needed to secure a manifold pressure may be as high a s 2200 rpm. However, once the required rprn has been established for an installation, any appreciable variation therefrom indicates malfunctioning. This may be because the low pitch stop of the propeller has not been properly set or because the carburetor or ignition system is not functioning properly. The accuracy of this check may be affected by the following variables : a. Wind: - Any appreciabIe air movement (5 miles per hour or more) will change the air load on the propeller blade when it is in the fixed pitch position. A head wind will increase the rpm obtainable with a given manifold pressure. A tail wind will decrease the rpm. atures: - The in atmospheric tend to cancel each her carburetor entrance atures tend to lower the rpm, but the propeller load is lightened because of the less dense

because of factors other pheric temperature, a I result as the power wil without a compensating the propeller load.

GROUND OPERATION-GROUND

CHECK

a,l

3. Magneto Check

*.

In performing the magneto che power absorbing characteristics of t peller in the low fixed pitch posit utilized. In switching to individu netos the cutting out of the opposite results in a lower rate of comb which gives the same effect as retar the spark advance. The drop in speed is a measure of the loss of attendant on this lower combustion r a t By comparing the rprn drop-off wit known standard the following are de mined : Proper timing of each magneto. b. General engine performance as e denced by smooth operation. c. Additional check of the proper co nection of the ignition leads. Any unusual roughness on either magnet is an indication of faulty ignition ca by plug fouling or by malfunctionin the live side of the ignition system. operator, should be very sensitive to engin roughness during this check. Lack of a drop-off may be an indication of fault grounding of one side of the ignition tem. Complete cutting out when switc to one magneto is definite evidence that its side of the ignition system is not func tioning. Excessive difference in rprn drop-off be tween the left and right positions can dicate a difference in timing betwee left and right magnetos. Inas Pratt & Whitney Air same spark advance ignition system, such corrected,

PWA. 01. 100

PRATT & WWITNEY AIRCRAFT 4. Fuel Pressure

FUEL PRESSURE

The fuel pump relief valve is set to give ormal i'light pressure a t the rpm of the ground check. Satisfactory fuel pressure indication a t this time is assurance that the fuel system is properly functioning.

than warm-up rpm must be kept to a minimum, especially if the aircraft is not headed into a good wind. It is essential that cylinder head temperatures never exceed the maximum specified for continuous operation, and it is desirable to keep them a t least 50F (30C) below the maximum continuous limit. It must be remembered that the head temperature indicator is connected to one cylinder only. This cylinder is selected on the basis of flight conditions, and often will not be the hottest cylinder during ground operation. 7. Idle Mixture Setting

5. Oil Pressure

O I L PRESSURE

ssure relief valve is set to give oil pressure a t the rpm of the eck and with a specified oil tematisfactory indication of oil presspecified temperature is assurequate oil pressure will be availoperating range. emperatures

Little, if any, cooling airflow is available on the ground, and operation at greater

Plug fouling difficulty is the inevitable result of failure to provide proper idle mixture setting. The tendency seems to be to adjust the idling mixture on the extremely rich side and to compensate for this by adjusting the throttle stop to a relatively high rpm for minimum idling. With a properly adjusted idle setting i t is possible to run the engine a t 450 rpm or even less for long periods with complete freedom from plug fouling. Such a setting will result in a minimum of plug fouling, exhaust smoking, and loading up ;and will pay dividends from ;he saving on the airplane brakes after landing and while taxiing. If the wind is not too strong, the check of the idle mixture setting can easily be performed during the ground check as follows : a. Close throttle b. Move mixture control toward idle cutoff and observe change of rpm c. Move mixture control back to automatic rich before engine cuts off As the mixture control lev ward idle cut-off, and off, one of two things tarily :

GENERAL OPERATING INSTRUCTIONS '

GROUND OPERATION

larly to ensure proper ground and idling.

(1) The engine speed may increase by as much as 200 or 300 rmp. An increase of not more than 20 rpm indicates proper mixture strength. A greater increase indicates that the mixture is too rich, as the engine acceIerates while the mixture leans out through best power,

NOTE: This check niust be performed in conditions. Any considerable wind, er head or tail, or across the ainpla affect the results.

Since the metering of the fuel by buretor in the idle range is not sated for altitude, i t follows that mixture setting correctly adjus field may not give satisfactor some other field of different the mixture is correctly set f of two fields, it will tend to be too the higher field; if correctly set higher field, i t will be too lower. In the case of an airplane whos flight schedule calls for sto fields of different altitudes, i t wil ally be found possible to make promise, intermediate idle mixt which will give satisfactory, if not results a t each of the various fie1 tendency of the mixture to be t o altitude, and hence to "torch" an the spark plugs, may be offset by speeds slightly higher than nor

(2) The engine speed may not increase, or may drop immediately. This indicates that the idle mixture is too lean, a s the fuel-air ratio has straightway leaned out beyond best power. The idle mixture should be set to give a mixture slightly richer than best power resulting in a 10 to 20 rpm rise after idle cutThis check should be performed with the idling speed set for 450 to 500 rpm. If the rpm is higher, correct idling adjustment will not result. In addition, the engine cylinder and oil temperatures should be a t a stabilized value representing the normal temperatures a t which the engine

ary adjustments made on new of time to obtain their

become soaked t is desirable

- IDLE -MIXTURE

8.

Taxiing Use a smooth flow of power or smoot changes of power during taxiing. and frequent "jazzing" sometimes feres with the operation of the a ing pump, with the result that backfi occur because of the low manifold pr sures just after a sharp closing of t throttle, while the engine is still tu a t a high rpm.

PWA. 01. 100

ATT i%r WHITNEY AIRCRAFT

COLD WEATHER OPERATION

operation of the aircraft engine require special preparautions a s compared to normal orization of the fuel behigh viscosity of the oil king speed with accompanys on the starter and batteries. ssories fail because of congealed cylinder priming washes the oil ston rings and cylinder walls causcoring of the cylinders, the bottom intake pipes PREHEATING

Approximately 0 F (-20 C) is the lowest temperature a t which engine starting can be accomplished using the standard priming system, accessories, and batteries. (Priming systems are being developed for some engines which permit consistent starts at -20 F (-29 C) or lower without the use of preheat.) Successful starting at- this temperature is contingent on adequate oil dilution, sufficient cranking speed, and proper technique. Inexperienced personnel will require considerable practice before achieving dependable starting a t 0 F (-20 C). Very capable personnel, using the conventional starting procedure, have made starts a t much lower ternperatures, occasionally as low as -35 F (-37 C). However, this practice is not recommended for general use because of the likelihood of hydraul-

icking and piston scuffing from the large amount of priming necessary, and the possible failure of the Wrication system or accessories. Therefore, the engine and installation should be preheated when ground temperatures are below 0 F (-20 C), unless the engine and installation are warm from recent operation, or have priming systems that .germit lower temperature starts. Adequate oil dilution is necessary to insure the required crankability for these improved priming systems. If oil dilution was not employed when the engine was previously shut down, external heat generally will be necessary a t temperatures below 40 F ( 5 C). Past experience, "stiffness" of the engine, and fluidity of the oil a t the Y-drain valve will be the best indications of the necessity for preheat. One suggested means of using engine covers and ground heating equipment for preheating the power plant is shown on page 93. In an installation of this kind, there are two regions which must be heated. The most important is the accessory compartment of the nacelle, the region a f t of the diaphragm or accessory cowl. In this compartment a sories, such as the star and the principal parts of the quent failures during or the accessory compartm inadequate.

!.

GENERAL OPE

GROUND OPERATION-COLD

WEATHER --

t

COWL F L A P d F R O M SOURCE O F HEATED A I R

I ,, DUCT TO 'SSORY SECTION

The se!c@ndregion in the nacelle which must be heated' is forward of t h e diaphragm. The power section of the engine is included in this space. The minimum time required for heating an engine is dependent upon the4capacity of t h e heater, the size of the engine, the outside air temperature, and t h e amount of oil dilution used prior t o the previous shutdown. For example, a heater delivering from 75,000 t o 90,000 Btu per hour will require about 30 minutes t o heat a Twin Wasp engine a t -30 t o -40 F (-35 t o -40 C) outside air temperature. With the same heater and same temperature, a minimum of 45 minutes wilI be required with a Double Wasp engine. The difference lies in the amount of metal in t h e engine which must receive heat. With strong ground winds or lower .outside temperatures, increased minimum heating time will be necessary. It is desirable t o heat the power plant for a longer period than the minimum t o ensure t h e easiest possible starting consistent with the time available for heating. When using preheat, care must be exercised not to burn the insulation of the ignition system, if the temperature of the preheat air exceeds 225 F (110 C ) .

The successful completion of preh:eating can best be determined by turning the propeller occasionally and noting engine "stiffness." This is particularly effective if no, or insufficient, oil dilution has been used. With extreme oil dilution, an engine will be free a t a temperature a s low a s -50 F (-45 C). Cylinder head temperatures, as indicated by t h e cockpit gage may not give the true cylinder head temperatures, if the preheated air happens to discharge directly in contact with the thermocouple. Checking the fluidity of the oil a t t h e Y-drain can be used as a supplementary check. This will also indicate the presence of ice which would block t h e flow of oil t o the engine. If the oil was drained after t h e previous flight, it should be preheated to a temperature of a t least 150 F (65 C) before being returned t o the aircraft oil system. Engine heating should be completed before the heated oil is returned. To ensure adequate voltage for the engine starter, warm and fully charged batteries should be installed and connected to the electric system a t this time. If the aircraft is equipped with a n auxiliary power plant, t h e latter should be started, or a ground source of auxiliary power connected to the aircraft electrical system.

PWA. 01. 100

PRATT & WHlTNEY AIRCRAFT

Normal starting procedure may be used in starting an engine properly preheated and with preheated or properly diluted oil. STARTING COLD ENGINES

In stai-ting a cold engine, ice may form on the spark plugs and on other parts of the combustion chamber after one or more unsuccessful attempts to start. The "ice" is a combination of fuel, oil and water, and will prevent the spark plugs from firing. Every effort must be made to "catch" the engine on the first starting attempt, since ice will form in the cylinder within a few seconds after it has fired and quit. If two or three starting attempts have resulted in nothing but a few feeble firing impulses, time can be saved and abuse to the batteries and starter can be avoided by removing a few front spark plugs to inspect for ice. If icing has occurred, the front spark plugs should be changed or the engine thoroughly heated before making further attempts to start. With the electric inertia or combination inertia-direct cranking starter types, be sure to energize the starter completely before engaging. Engine priming is dependent upon the type of priming system and the amount of preheat which have been used, and on other factors that have been discussed elsewhere under "Priming". With sufficient preheat a normal warm engine start may be made. With pressure injection carburetors, after engaging the starter,.move the mixture control slowly out of idle cut-off as soon as the engine is firing on the prime, but be prepared to return it to idle cut-off quickly if the engine quits. Particularly with a stiff engine, it is very easy to get the mixture too rich for running and the engine will not pick up speed. Therefore, brief periods during which the mixture control is moved back to the idle cut-off may be used to advantage until engine speed increases enough so that airflow is sufficient to match the fuel flow. Caution must be used to avoid introducing excessive amounts of raw fuel to the lower in-

take pipes. Frequent instances of hydraulicking will occur if this precaution is overlooked. Except with a warm engine, a small discharge of fuel from the supercharger drain is normal while attempting a start. With float carburetors the mixture control will be positioned in full rich or automatic rich when starting, and no further manipulation of this control will be necessary. In cold weather, engine operation immediately after starting is frequently rough, with back-firing and after-firing. This is due principally to a lean carburetor idling mixture and to reduced vaporization of the fuel. Fouled or "iced" spark plugs will also produce the same effects. As a corrective, turn on the carburetor air heat as soon as the engine is free of backfiring. This will increase the fuel-air ratio in the idling range and will improve vaporization of the fuel. If the engine is equipped with a blower-rim or blower-throat priming system, the intermittent use of priming will assist in preventing backfiring until sufficient heat is available and carburetor air heat can be used. After the engine has warmed up operation will be satisfactory with cold air. WARM-UP

No attempt must be made to heat the engine rapidly after starting by closing the cowl flaps. As in normal weather, keep the cowl flaps fully open a t all times on the ground. Cylinder temperatures are not as important as oil temperatures, which represent the temperature of the internal engine parts. Cylinder temperatures will rise quite rapidly to temperatures only slightly below normal. Consequently the principal problem in warm-up is to increase the oil temperature. Until the oil temperature increases to 140 F (60 C), keep manually controlled oil cooler flaps fully closed. In the case of engines that have had the 011 diluted with fuel, i t is preferable to allow the oil temperature to rise considerably above 140 F (60 C ) , and increase the engine speed during warm-up to 1200-1400 rpm. Very little gasoline is driven out of the circulating oil below this oil temperature and engine speed. As much of


GROUND OPERATION-COLD

the diluting fuel as possible should be eliminated during ground warm-up. With heavy dilution, the gasoline is evaporated very rapidly during the take-off with a tendency to discharge oil from the breathers, unless eliminated during warm-up.

WEATHER

minute) into the oil inlet line, usually a t the Y-drain valve. Mixing of the oil and gasoline is accomplished within the engine by the mechanical agitation of the pressure pump, moving internal parts, and the scavenge pumps.

During warm-up, hydraulically operated propellers should be "exercised" several times to make sure that the propeller mechanism is free of congealed oil. So that the propeller governor will operate, briefly bring engine rpm up to 1500 or 1600.

FLIGHT OPERATION Flight operation in cold weather is hardly more difficult than in normal weather. After starting and ground operating troubles have been overcome, and a successful take-off has been made, flight operation is practically the same as under standard conditions.

OIL DILUTION Oil systems may be completely drained after the engine is shut-down, and the oil heated and returned to the system a t the time for the next start. Some aircraft are equipped with electric immersion heaters in the oil tank. Used continuously or in intermittent periods the contents of the tank, with the possible exception of the base of the oil tank hopper, will be kept in a fluid condition. Before starting, the engine and accessory compartment must be thoroughly heated by means of ground heaters. I

More generally, an "oil dilution system" is used to ensure fluidity of the oil after the installation has cooled. Oil dilution consists of introducing raw fuel into the oil to reduce the viscosity of the oil. Because of substantial differences in the specific gravity and viscosity o f gasoline as compared to aviation oil, there is very little tendency for them to mix when introduced into a common line or tank. However, if the two fluids, in any proportion, are forcibly brought together by some means of mechanical agitation, a very permanent mixture is produced. Diluted oil will not separate if allowed to stand, provided the oil and gasoline have been thoroughly mixed. The Army type of oil system is designed for effective use of oil dilution. Gasoline is introduced a t a slow rate (1% to 2 quarts per

a h v e ilkastmtion shows a typical Army ail ~ysternrepresenting conditions in the lines 8nd tank a few s.eonds after dilutiofi has been s-briml

Gamline i~ led from a pressure take-off conneetian; passes th.rough a solenoid operated valve; and enters the oil inlet line, usually at it wnnwtion on the Y-drain valve body. The mlenoid valve ia operated by a switch in the wkpit.

An e~i$entialpart of the oil system designed for the use of dilution is the oil tank hopper. The hopper is a 1 to 2% gallon capacity cylindrical ves,sel mounted near the center of the main oil tank. The hopper serves to r e d m the amount d?oil in eircubtion by separating the main r w r v e sii supply from the eirewhting oil. With only a small arnannt of oil in circulation, warm-up is rapid and dilution in the higher pementage~~ (2Wb to 35%) is made possible. Without a hopper, the entire oil supply would have to be diluted, requiring a very large expansion space. By reducing the amount of oil ta be diluted and also By increasing the rate of warn-up, the hopper reduces tbe length of time Bu4in.g which the engine must operate with diluted oil a . f i ~ rsstorWng. -. - ..p The.illa&r.ation on the following gage s ows, the Army oil system after dilution is complete. All oil lines and the hop-per are diluted to the s a w percentage, wkieh is also true of a!l oil ih.

rn

PWA. 01. 100

PRATT & WHITNEY AIRCRAFT the engine. Some of the diluted oil will also find its way outside the hopper and will collect on top of the undiluted oil in the main tank. The amount of oil dilution necessary is dependent upon the expected temperature to be protected against. Oil dilution is recommended when the expecteu starting temperature is 40 F ( 5 C ) or lower. The elapsed time required to dilute to a desireci percentage depends on tvr7o factors ; the rate of flow of gasoline into the oil, and the amount of oil to be diluted. This fact explains the difference in time interval required for various aircraft. I t will be necessary for the operator to consult the manufacturer's specific instructions for the aircraft concerned. Oil dilution is accomplished just before shutting down the engine. Before beginning oil dilution the engine should be idled, if necessary, to reduce cylinder head temperatures to a maximum of 300 F (150 C), and allow the oil temperature to fall below 120 F (50 C). While idling the engine, close the oil dilution switch for the required time. Keep the switch closed while stopping the engine, and until i t has stopped turning over. If i t is necessary to service the oil tank, this should be done prior to starting dilution.

I n extreme weather conditions, it is necessary to dilute the oil going into the Hydromatic propeller by "exercising" the propeller governor control during the latter part of the dilution time interval. This must be done when the engine is briefly brought up to 1500 o r 1600 rpm so that the governor will operate. Oil operated supercharger controls require the same treatment under these weather conditions. The introduction of gasoline into the lubricating oil will loosen carbon and sludge deposits within the oil system. This carbon and sludge is carried to the engine oil screen and collects there in quantities sufficient to cause the collapse of the screen. Consequently, within a n hour or two after the dilution is first used in the season, the screen must be removed for inspection and cleaning. This inspection must be repeated a t short intervals until sludge and carbon no longer collect. Should this precaution be overlooked, collapsed screens are certain to result. Trouble with dislodged carbon may also be encountered if dilution is not used almost daily following its first use in the season. Some northern operators use small amounts of oil dilution throughout the warmer weather to keep the oil system cleared of carbon and sludge.

Dilution Completed

FLIGHT OPERATION

-

-

-cg

1 Considerations

..

--

a r t of the take-off run the engine, time since starting, is encounterns resulting from the use of high

--

power. The considerations governin proper balancing of all operating factor now apply.

used, and there should be e-off rating, the"type test r s a t this output, or the of over 1000 take-offs.

ions permit, it is

ligh pow,er and 1 97

The pressure is applied by adva regulators will prevent exceeding allowable limits, but on all other aircraft i t will be necessary to give the manifold pressure gage sufficient attention to prevent "overboosting". Engines equipped with multiple speed or multiple stage superchargers usually require that the lowest impeller speed or stage be used a t take-off in order to prevent excessive temperature rise through the supercharger. Exceptions to this rule may be allowed on specific engines. For complete information consult the applicable specific operating instruction.

that no more

he supercharger. Howfor full take-off power a t any

On installations with fixed or two-position propellers, the limit of throttle advance is defined by maximum manifold pressure and/or maximum rpm. At the one extreme of low altitude combined with low free air temperature, limiting manifold pressure may be reached a t less than limiting rpm. At the other extreme of high altitude and/or temperature, take-off rpm may be obtained a t less than take-off manifold pressure. It is essential that both instruments receive the necessary attention a t this time. 4. Mixture

-Fuel-Air Ratio

FLIGHT OPERATION-TAKE-OFF

GENERAL OPERATING INSTRUCTIONS --

.

head temperature caused by lack of cooling airflow requires that the mixture be considerably enriched during the take-off. To provide this mixture use automatic rich in the case of autdmatic carburetors or rich in the case of manually controlled carburetors. As a result of the previous checks of engine operation and fuel pressure the pilot will have the assurance that the required fuelair ratio will be obtained. 5. Charge Temperature Temperature

--

-.

6. Cylinder Temperature

-Carburetor Air

a. Rlaximum Temperature . .

-

Conditions may a r k when i t is neeewry to use carburetor heat to eliminate ice b fore, md even during, the take&. When ice has formed during ground opepation a short run to rats rpm with the carburetor temperature control in "Full Hot" will sufiRce to clear out the induction system. The comtro~then &6uId be returned to "Cold" befom beginning the take-off.

In the m Q l y possible event that hesit is needed during take-&, w&ient athnticr~ must be giyen to the temperature indicator and tho control to prevent the carbwator air temperature exceeding the r m m mended limits set for any specific engine model. Obvi-o~sly,when the free air temperature & near or a h v e the m m m d e $ limit i t is not p d i b l e to maintain: this limit.

The high pressure loads of take-on power require that the cylinder head strength not be reduced as the result of excessive temperatures. I t can be anticipated that the head temperatures will increase 45-55 F (25-30 C) during the run as the average cooling airflow is insufficient to cool the engine at this output on a continuous basis. Therefore, before the run is started the head temperatures must be sufficiently low to prevent this rise from exceeding limits. Even in the coldest weather a cowl flap opening greater than the closed position is essential. The ignition narness especially needs more cooling airflow tharl closed flaps will provide during the take-off.

PWA. 01. 100

PRATT & WHITNEY AIRCRAFT b. Minimum Temperatures

'

'1 9,Y

.r

\r"b7.iq

There is no fixed minimum cylinder head temperature. Satisfactory engine operation has been obtained with head temperature as low a s 76 F (26 C) ; and i t is not known if this is the low limit. Warmer temperatures are recommended a s they usually assist in fuel vaporization and distribution and, hence, smoother running. 250 F (120 C) has been suggested as a minimum based on general considerations of overhall engine conditions. It should not be eonstrued from any such instruckion t h a t i t is necessary to employ Unusual means t o bring thme heads up to any specific limiting minimum temperature. The oil t e m p e r a t u e is the best gage of the over-all engine conditions. 7.

m

t o critical bearings for cooling purposes. No take-off should be made until t h e oil has warmed up a t least t o this point and 140 F (60 C) is preferable. The maximum temperature prior to take-off should allow for the rise which inevitably follows, similar t o the rise in head temperatures. This increase will vary with different installations, depending on t h e capacity of each system, whether or not all the oil contained in the system is used continuously, and the type of temperature reguIation. The maximum limit must not be exceeded as high temperatures can result in foaming and discharge through the breathers, a s well as the natural reduction in lubrit n g qualities. This is especially true while operating under conditions of take>ff power and temperature. As take-off power and speed a r e the conditions of the greatest need for lubrication, ithe functioning of the lubrication system,

e carefully checked.

Lubricatio

100 F (40 C) is t h e a b s w e minimum of t h e oil inlet temperature h g e which will ensure proper lubrication and flow of oil

No take-off shou1.d be performed without assurance of proper ignition system functioning. While less than perfect ignition operation can be tolerated under conditions of low engine power and speed, the high pressures and high intake temperatures encountered during take-off require t h a t t h e entire system be functioning in order that the engine may maintain maximum output without detonation. The magneto check at approximately 65% power has been found t o satisfy this requirement.

-xn-

GENERAL OPERATING INSTRUCTIONS PREPARATION FOR TAKE-OFF CHECK-OFF LIST

-

LIGHT OPERATION-TRANSITION

TO CLIMB

TRANSITION FROM TAKE-OFF TO CLIMB

As soon a s the field and surrounding obstacles are cleared, reduce power a t least to Normal Rated. With constant speed (variable pitch) propellers the reduction should be accomplished in steps as follows:

A properly prepared and consistently followed check-off list will ensure that all necessary steps are taken while preparing for take-off. The take-off run with its demands for

eck will give the pilot a clear rning items which otherwise is a suggested power plant

- Full

open

- Set for take-off

- On suitable tank* i r Heat - Cold (except a s discussed) i r Filter - AS required - Low, Neutral, or Off

1. Retard throttle to reduce manifold pressur to about 2 in. Hg below that for Norma Rated power (with fixed part throttle, mani fold pressure will rise as rpm is reduced).

2. Retard the rpm control to Normal Rated rpm.

If a further reduction in power is desired, proceed as follows: 1. Lower manifold pressure by 2 in. Hg. 2. Lower engine speed by 200 rpm. Continue in successive alternate steps until the desired engine speed is reached, finally adjusting the throttle to the desired manifold pressure.

Oil Pressure Fuel Pressure Cylinder Temperature - Adjusted for take-off - Adjusted for take-off c aircraft would

This is not to be construed to mean that the throttle should never be advanced with a low rpm. The engine is not affected by the position of the throttle-it.-is affected by the manifold pressure resulting from the throttle position.

-

PWA. 01. 100

PRATT & WWITNEU AIRCRAFT

I

I

0

TRUE AIRSPEED

I I

-MPH

Fig. 29 - Brake Horsepower vs True Airspeed

POWER her than specifying maximum limits of engine manufacturer cannot preite outputs to be used for the variitions of flight. There are no direct s to the questions: "What is cruise power) ?", and "What is climb (power) ?" The pilot has available any power from the minimum necessary to sustain the aircraft dp to Normal Rated. What he should use is governed by what quality of service he wishes to emphasize : performance, or economy and dur1. Aircraft Performance vs. Power

The relation between aircraft speed and engine brake horsepower is illustrated by Fig. 29. The significant information to be obtained from this curve is the low return in increased airspeed for a given increase in power a t high output as compared to medium outputs. The upper portion of the curve follows closely the relationship given by the equation:

where V,, V, = two differentairspeeds Bhp,, Bhp, = the brake horsepowers necessary to obtain the correspond-

and

ing speeds.

The equation, it will be noted, is similar to the propeller load formula, and in nonmathematical language simply means that 8 times the power is required to double the air speed, and 27 times the power to triple it. Stated another way: if the bhp is increased in increments of 10% Normal Rated power, an airplane which develops an airspeed of 220 mph a t 4021 of power will realize a gain of 20 mph as the bhp is advanced to 50% of power; but will increase its speed by only 10 mph (i.e., from 289 to 299 mph) as the bhp is increased from 9021 to 100% of Normal Rated power. 2. Performance vs. FueI Consumption Perhaps of even greater interest is an illustration of the reIationship between airspeed and miles per gallon.


FLIGHT OPERATION-POWER

SELECTION

I TRUE AIRSPEED-

MPH

Fig. 30 -Miles Per Gallon vs True Airspeed

It should be obvious from this curve that economy of fuel cannbt be realized by operating a t a high percentage of power. The decrease in miles pev gallon as speed is increased is the result of the rapid increase in power required, combined with the fuel enrichment accompanying high power operation.

This is another way of stating that an engine can be used for a certain number of power-hours before requiring a n overhaul. These power-hours can be used up in a short time by high power operation or they can be stretched by reducing the percentage of rated output used continuously. 4. Power Selection

Performance vs. Durability

The complete picture of the effect on economy and durability of varying airplane performance can be shown by combining Figs. 29, 30 and 31.

The durability of the engine as measured in the number of hours between overhauls is also a factor which varies with the amount of power used, and so can be shown as varying with airspeed.

The operator has the entire range of performance to select from and in many cases

TRUE AIRSPEED

-MPH

Fig. 3 ? - burabiliQ vs True Airpeed

I

PWA. 01. 100


/

fig. 32 - Brake Horsepower, Fuel Economy and Durability vs True Airspeed

el economy and engine dvantageous to tinuously. Miliin a manner dicnecessity ; durasecondary considera' , after a thorough ana s more profitable to : a t high speed, s earnirig more revenue, even though ine operation costs per mile are in3sed.

BMEP

- CONTROL

When such considerations are not present it is suggested that power be governed a s 1. Climb

wer.

)f

2. C r u i s ~ .,..,st pow Lnce needs, not rmal Ra ted pow

The accun.....,,, experience of severaI yea.rs has iindicate(1 that operation within the se limits will Iresult irI the optimum --_ -r - balance 01 aircraft perlorman erall economy of operation. . L -

reflight Preparation Prevents Perplexed Pilots ? power to be used, the lose the engine speed for

tion Propellers I ~ installed, L . only If a fixed ~ I L C L I~ L - u ~ ~ I is one combination of manifold pressure and

GENERAL OPERATING INSTRUCTIONS rpm will give the required power a t a given altitude and airspeed. This is also true of either the high or low pitch position of a two-position propeller. With a two-position propeller the requirements of proper rpm and manifold pressure balance will be best met by operating in the high rprn position during take-off and climb and in the low rpm position during the cruising portion of the flight. Otherwise there is no opportunity to exercise bmep control and establish the optimum combination of manifold pressure and rpm. 2. Variable Pitch (Constant Speed) Propellers

With variable pitch (constant speed) propellers an infinite number of rprn and manifold pressure combinations may be selected to deliver a given power, bearing in mind that, a t constant power, as rprn is increased, manifold pressure must be correspondingly decreased, and vice versa. For example: engine "A" a t sea level will develop 650 bhp a t 2600 rprn and 28.75 in. Hg manifold pressure, a t 1450 rprn and 37.0 in. Hg, or a t an infinite number of intermediate combinations. In selecting a particular combination of rpm and manifold pressure to obtain a desired power, the operator must regulate the balance between the loads and stresses due to engine speed, on the one hand, and those attributable to cylinder pressure on the other. Bmep accordingly becomes one of the principal factors, not only in the selection of rprn and manifold pressure, but also in determining economy of operation and engine durability. 3. Engine Speed vs. Cylinder Pressure a t Constant Power

In the case of an engine operating a t constant power, but a t different rprn and inversely varying manifold pressures, the effects of speed (rprn) and cylinder pressure (bmep) on various characteristics of its performance may be summarized as follows : a. Friction: - Reduced by a decrease in rpm. The decrease of sliding and rolling speed more than offsets the corresponding increase in pressure friction.

FLIGHT OPERATION-BMEP

CONTROL

b. Reciprocating and Centrifugal Loads : - Reduced by a decrease in rpm. Pressure opposes and cushions these loads which predominate in determining the stresses on master road bearings. c. Detonation: - Reduced by a decrease in rpm. Inasmuch as the tendency of a fuel to detonate is determined principally by its temperature sensitivity, a low impeller rpm with the attendant low temperature rise through the supercharger is favored. The heat rise varies chiefly with impeller speed and is relatively unaffected by power. Secondarily, the tendency of a fuel to detonate is determined by its pressure sensitivity. Excessive manifold pressures combined with high mixture temperatures can therefore also lead to detonation. d. Fuel Consumption: - Reduced by a decrease in rpm. As explained above, an increase in rprn is accompanied by

MINIMUM

RPM

MAXIMUM

Fig. 33 -Fuel Consumption vs Rpm at Constant Bhp

an increase in friction. Accordingly, if a constant brake horsepower is to be delivered to the propeller shaft, i t follows that the cylinders must develop more power (ihp), and this additional cylinder power requires extra fuel. e. Cylinder Cooling: - Improved by a decrease in rpm. The reduced cylinder horsepower (ihp) resulting from low rprn operation will be reflected in lower cylinder head temperatures.


PWA. 01. 100

f . Oil Cooling: - Improved by a decrease in rpm. Low engine speed is the most effective way to control the tendency of the oil to overheat, since a low rprn reduces the power necessary to overcome friction and, hence, the amount of heat to be carried away by the oil.

ever, because of structural and detonation considerations. These limits can be described in terms of bmep, and may be illustrated as follows : 4. Maximum Cylinder Pressure Limit

A line joining Ovi power and rprn with 100:4. normal rated power and rprn defines the minimum recommended engine speed for normal rich mixture operation. Higher bmep should not be used unless there is a reason proved by analysis for

g. Altitude Performance: - Above the

critical altitude for a given combination of manifold pressure and rpm, manifold pressure decreases, and with it. the ennine's power output. Under these circumstances the desired power can be maintained only by increasing engine speed, thereby raising t h e pumping capacity of the cylinders and increasing the pressure rise through the supercharger. h. Vibration: - It has been the practice to select engine and rprn combinations which appear t o give acceptable cockpit and cabin comfort. Engine suspension systems of earlier types tended to favor a bmep lower than would be selected from considerations of engine efficiency. Present developments of engine suspensions are such that, in general, there is little if any limitation to the use of the full rpm range and meeds a s low a s 1200 r ~ m may be entirely practical.

certain P r a t t & Whitney Aircraft enodels now in current use. For in-

Recommended Maximum Lean

It may be concluded from the foregoing discussion t h a t i t is generally preferable t o operate the engine on the side of high manifold and cylinder pressures and low rpm. Cylinder pressure must be limited, how-

50

I00


Fig. 34 - Brake Horsepower vs Rpm

- (Bmep)


FLIGHT OPERATION-BMEP CONTROL

a

g

80

a

W n

I-

2 so _I

i 0

40

B C W

ga 20 - oO

10

20

30

40

50

60

70

80

90

1

100


Fig. 35 -Brake Horsepower vs Rpm

5. Maximum Rpm Limit The high rpm limit of operation can be conveniently defined by a propeller load curve passing through Normal Rated power and rpm. Operation a t engine speeds lower than propeller load will insure that the loads set up by high rpm are never excessive and will be sufficiently cushioned by pressure. 6. Desired Rpm Range Combi6ing the pressure and rpm limits, the area of best operation would appear a s shown in Fig. 36. It will be noticed that a considerable range of power and rprn combinations is still available a t medium percentages of the normal rating. Exact definition of the optimum combinations for operating in this area must come from the aircraft

- (Propeller Load) manufacturer. On practically all counts the engine will wish to run a t t h e maximum bmep. The propeller prefers an engine speed more nearly at propeller load. The airplane by itself will wish t o be flown somewhere in between depending upon its own characteristics. The answer can only be obtained by combining the three units together. If maximum miles per gallon are desired, the engine will say, "Operate me a t the lowest brake specific fuel consumption." This may result in a propeller speed so low that the blade angle will increase to the point where i t is "paddling" the air and providing Iittle thrust. The engine will feel very satisfied a s i t will be using the Ieast gallons per hour, but the airplane will not go very f a r with a given amount of fuel because of the inefficiency of the propeller.

OF EXPERIENCE AS DEFINING THC RANGE OF GREATEST ECONOMY AND DURABILITY-.

PERCENT OF

NORMAL

RATED R P M

Suggested Maximum Bmep

,

can be

atisfied if the power and rpm

the engine involves infig. 38 - Brake Horsepower v

these conditions is deter be taken into account:

ariations : - Head o vor a deviation fr ro-wind altitude. ye opera --:Ll-

.


8. Induction System Icing Ice formation within the induction system is often a potential hazard of engine operation. It is difficult to approximate the frequency of this condition. The more advanced cases of icing, which cause serious malfunctioning of the engine, are comparatively few in number, but they cause a great deal of concern, and should be absolutely eliminated. ed in the induction system processes which are

AIRSCOOP E

SCREEN IMPACT TUBES

BOOST VENTURI

TI

ICING BELOW THROT NOT LIKELY I F FUEL D/SCHARGED HERE

THROAT VANES

is that formed from

FLIGHT OPERATION-ICING ing after it is introduced in stream. This ice probably occurs frequently in actual operation be it may form a t carburetor air tempe tures considerably above 32 F (0 C). Most of the heat necessary to evaporate the fuel is taken from the air, which causes i t to drop in temperature Fuel ice may affect airflow by block ing off the supercharger entrance (blower throat), affect fuel-air ratio by interfering with the fueI affect mixture distribution or quant of mixture to individual cylinders upsetting the fuel flow a t the zle distributor, or air flow dis a t the supercharger entranc certain conditions of high humidity, this ice may form with carburetor air temperatures as high as 80 F (25 C) Tendency for fuel ice to form is greater with float type carburetors and to a lesser extent with pressure carburetors having the X-bar fueI d tributor. The newer type spi distributor developed by P r a t t & Whitney Aircraft for the pressure injection carburetor has practically eliminated icing of this type. Indication of icing conditions in the induction system in the order of probable perception to the operator are as follows: a. Decrease in manifold pressure-due to restriction of induction passages, with consequent loss of power. b. Changes in fuel-air ratlio; the mixture becoming either richer or leaner. This may become serious before there is a appreciable change in manifold pre sure. , c. Sticking of the throttle valve. d. Icing indicator instruments

-

(if in-

tion depend upon the manner in which the is discharged into the induction syste

.

ATT & WHITNEY AIRCRAFT and Whitney Aircraft engines with float carburetors and the Twin Wasp C series). When the fuel is injected immediately downstream from the carburetor a considerable volume is available in which fuel evaporation ice can form. Turning vanes offer a convenient surface for the ice to cling to and complete blocking of the passage can quickly occur. With this type of equip-

drop due to evaporation ce the temperature in this elow freezing. 100 F (40 C) carburetor air temperature should be used when possible ice formation conditions are present. As the evention of evaporation ice requires ree of heat the other possibilities are taken care of adely. Do not apply greater pre-heat

PWA. 01. 100 of 100 F (56 C) or more will insure the abiIity to remove ice already formed. Some engines in this category require definite restrictions in the use of preheat in order to prevent detonation. This is especialIy true when using the high impeller ratio of a single-stage, two-speed engine. For more definite information regarding individual engines refer to the applicable specific operating instructions. Removal of ice already formed is best accomplished by use of full carburetor heat. If the preheat capacity is sufficient and the remedial action has not been delayed, i t is a matter of seconds until the ice is removed. The preheat capacity can be increased by applying more power and by closing the cowl flaps. If the ice formation is alIowed to progress to a critical extent the loss of power may

uretor air temperature represents ossible loss of engine performance d excessive carburetor heat can Iead USE HEAT FOR PREVEN'TION IT MAY N O T BE THERE FOR CURE

heat. It is possiwhen using 6ccoId"air, for the coned moisture to pass through the temperature is elevated to 32-80 F (0-25C) this moisture will be changed first from a frozen to a liquid condition and, when fuel evaporation takes place, will return to the frozen state in the induction passages. Under such condi(40 C) preheat air temperature controI injected a t or downstream from ner injection.) With this equipment fuel evaporation ice is no longer a criticaI factor. However, while a large preheat capacity is not needed to prevent ice formation, the instant availability of a temperature increase

make it impossible to generate sufficient heat to d e a r the engine. Because of this possibility it is imperative that the crew be alert to possible icing conditions and take remedial action while i t can still be effective. Operation of the engine a t high bmep (large throttle opening and low rpm) may assist in minimizing the possibility of ice by reducing this obstruction across the airflow. Also, a change of altitude may result in finding ice-free conditions. There is record of some operators clearing ice by causing the engine rigorous procedure should be resort only in extreme emergency and it i tant to insure that the carb trol is in the "cold" position "off" so as to prevent darnag

wered airspeed. An airplane towing a level flight is subjecting its engines me conditions as in an unencumbered

climb a t t h e same indicated airspeed. The higher powers used for climb combined with reduced cooling airflow call for closer attention t o cylinder and oil cooling.

PRATT & WHITNEY AIRCRAFT SPEED AND PRESSURE - RPM AND MANIFOLD PRESSURE

elect rpm to give power within limits shown Fig. 36. Greater than Normal Rated bmep ould be used only when i t is known that finite advantages are obtained. If the climb is d above the critical altitude for the nd rpm selected, increase the engine rpm for each inch reduction of maniURE

-FUEGAIR

PWA. 01. 100

latch cover of the mixture control can be felt through the cockpit control, and proper positioning positively assured. As new developments take place different nomenclatures may be used to denote proper mixtures. Consult the applicable specific operating instructions. 2. Non-Automatic Carburetors: Rich mixture is obtained on non-automatic carburetors in the full forward position of the control. Above 5000 feet it may be necessary to retard the control to compensate for the effect of altitude on the uncompensated carburetor. This retarding should be limited to the amount required to maintain smooth operation. CHARGE TEMPERATURE AIR TEMPERATURE

-CARBURETOR

RATIO

misture is used in climb to assist the airflow in maintaining the desired cylteract any tendbecause of the ercharger when the percentage of 1. Automatic Carburetors: On current engines provided with automatic carburetors, rich mixture is provided by positioning the control in automatic rich. An indent on the

Any necessary use of preheat during the climb must be closely controlled so as not to increase to a critical value the high mixture temperature resulting from high rpm. If detonation is encountered during the climb, the most effective corrective action is to reduce manifold pressure and rpm. The rpm is reduced to lower the impeller speed, thus reducing the supercharger heat rise. The lowering of manifold pressure is necessary in order to stay within bmep limits. The carburetor entrance temperature limit for single-stage, single-speed engines a t n rated rpm, or greater, is usually 100 F (4 When higher entrance temperatu quired to eliminate ice already f manifold pressure and rpm so charge temperature will not be exce

a s possible c smooth operation non-automatic ca

Two limiting cylinder head temperatures are

LUBRICATION

- "..---

re is allowed for a limited

high stresses on the head

weakens the heads.

r these

eratures with-

smbustic roper lii

..

.

provided, the oil cooler control maintain the desired tempera control, whether manual or aut give the necessary temperature the following steps : 1. Reduce rpm and manifold pres The rpm reduction reduces the

.ircraft sary coc

engine friction. The rnanif duction may be necessary

aximum Cont rna~imui continu 7.7.

g climb

..

2. Increase airspee

PWA. 01. 100

RATT & WHITNEY AIRCRAFT

TRANSITION FROM CLIMB TO LEVEL FLIGHT I-

-

PEED AND PRESSURE PRESSURE ,

-

- RPM AND

$...

I

Delay closing cowl flaps or leaning mixture until the cylinders have cooled down below 400 F (200 C). This delay serves not only to cool the cylinders after the climb but to bring to desirable temperatures the entire powerplant. Close the cowl flaps gradually as the speed increases.

MIXTURE

- FUEL-AIR

RATIO

flight power by first reducing

--

JIC*

- --

-'

Change to cruising mixture after the engine been attained. The mixtu prevent the combination a warm engine. An e engine can resul will prevent the

FLIGHT OPERATI

GENERAL OPERATING INSTRUCTlONS 1

LEVEL FLIGHT Approximately 95% of flight time is spent in this engine operating condition and, therefore, has the greatest influence on engine

POWER: The lowe

ive the desired aircraft performance within the following limits:

Recommended All Normal Operation - Maximum Cruise Rating All Other

PWA. 01. 100

PRATT & WWITNEY AIRCRAFT SPEED AND PRESSURE MANIFOLD PRESSURE

- RPM AND

realized. It is intended that cruising lean mixtures be used under favorable conditions of operation. When it is not possible to maintain cylinder temperatures within maximum continuous limits the mixture must be returned t o the rich or automatic rich position. Many pilots have noticed a change of power while holding constant manifold pressure and rpm when reducing the mixture strength t o lean values. The loss of power is directly evident, if a torquemeter is installed, or it may be noticed by a slight reduction in airspeed. This drop in output is the normal result of reducing mixture strength below best power 7

7

-

Obtain desired power within bmep limits as own on Fig. 36. Where aircraft manufacrer has provided complete cruise control inormation, use his data for precise control.

- FUEL-AIR RATIO

Fig. 39

on in excess of Normal Rated s t be with rich or automatic rich mixuous operation. For powers beRated, Pratt & Whitney Aircraft limits for power, rpm and bmep lean mixture. While these recom~ t sare not mandatory the use of a t greater powers should only be the basis of analysis as to the to be obtained. The use of lean or ean mixture a t any time must be refavorable conditions of cylinder and

-Power vs Fuel-Air Ratio

fuel-air ratio and does not mean that the economy of lean mixture operation is being lost. If the operator has available operating curves covering automatic rich and automatic lean operation, the extra manifold pressure necessary to obtain the desired performance can be readiIy determined. When such information is not available the throttle should be advanced to regain the airspeed lost when the mixture was leaned. 1. Automatic Carburetors

The use of lean or automatic lean mixture under cruising conditions is essential if the potential fuel economy of the engine is to be

Current engines, equipped with automatic carburetors, have the lean mixture position indicated as Automatic Lean. When the

FLIGHT OPERATION-LEVEL


FLIGHT

REOUCTION I N FUEL-AIR RATIOIN RANGE SHOWN IS NO? APPRECIABLE PRACTICAL MANUAL LEANING Rf OUIRES CREATER ANGLE OF MANUAL CONTROL

M/XTURE CONTROL POSITION

Fig. 40 - Fuel-Air Ratio vs Mixture Control Position

mixture strength is to be reduced the control should be retarded by placing it in the automatic lean position. An automatic lean indent is provided that can be felt during the control motion. (For carburetors with normal and rich positions, refer to Page 31.) The question is raised in many sources as to the possible advantages of manual leaning between, automatic lean and idle cut-off. On carburetors of previous and current manufacture, this is not a procedure to be attempted except under conditions allowing the closest attention to this operation and in installations provided with instrumentation for the purpose. Furthermore, the relation between mixture control travel aria fuel-air ratio is extremely sensitive and such that accurate control is difficult. It will be seen that a sudden drop in mixture strength occurs with a very slight motion of the control. Under all normal conditions it is, therefore, recommended that lean mixture operation be performed with the control placed in the automatic lean indent. As new developments take place, individual models of carburetors may be provided with mixture regulation allowing manual leaning and a different nomenclature may be used to distinguish proper mixture positions. For exact information consult the applicable specific operating instruction.

2. Non-Automatic Carburetore a. With manually controlled two-position, or fixed position propellers: Referring to Fig. 39, i t will be noted that, a s the rich mixture used in climb is leaned out, the power developed increases until the rich best power setting is reached ; remains virtually constant a t a maximum between rich best power and lean best power; and decreases rapidly as the mixture is leaned out beyond lean best power. When a propeller is used which can be held in one pitch position, any change in power is accompanied by a corresponding change in rpm. This relationship is made use of in setting the mixture for level flight as follows: (1) Fix throttle to the desired rpm and manifold pressure and adjust carburetor air temperature for continuous operation. (2) Lean mixture until rpm reaches a maximum value. This is rich beet power and is the full extent to which the engine may be leaned when operating between 65-75% power. In leaning to this condition it will be normal to go slightly beyorid the point where maximum rpm is first noted. The control must be returned to this point, making sure that control link age backlash is accounted for.

.

& WHITNEY AIRCRAFT

r the mixture can an best power by ntrol beyond rich the rprn starts to lean beyond a 10 rom maximum rpm.

h the values specified in the specific operating instructions for the engine. (2) No Instrumentation With no instrumentation, manual leaning cannot be done safely as there is no reaction from engine or instrument to advise the operator a t safe conditions are being mainined. Manual leaning is permitted to the extent required to smooth ine roughness resulting from ect of altitude on the uncomed carburetor. When manual mixture adjustment is used it ust be remembered that the control position yives the correct fuel-air ratio only for the coniitions uncler which the adjustment is made, lamely: pourer, rpm, carburetor air tempera;we, and altitude. If any of tllese conditions are changed, the mixture must be reset. There is no definite mixture control position corresponding to any specific combination of manifold pressure and rpm, but the following will serve as a formula to enable the operator to maintain a proper mixture strength : re Control (En-

CHARGE AIR TEM

With Wasp J r , Wasp, Twin Wasp J net engines, or other Pratt & Whitney Aircraft engines using float-type carburetors, it is desirable to maintain a constant 90 F (30 C) carburetor air temperature to ensure freedom from icing and to assist mixture distribution.

On all engines it will be necessary, under ice formation cooditions, to increase the carburetor entrance temperature either to eliminate ice already formed or to maintain a carburetor air temperature that will ensure freedom from ice formation. In such circumstances the immediate concern is freedom from ice, but it should be realized that as an indirect result the mixture charge temperature is being increased. The maximum limit, usually 100 F (40 C) , must be observed. Refer to the applicable specific operating instructions for the engine i tion. The use of a 100 F (40 C) carbur formation a t all times. formed it may be ne preheat capacity to cl

FLIGHT OPERATION-LEVEL

GENERAL OPERATING INSTRUCTIONS There is no minimum carburetor air temperature. Full cold operation in sub-zero conditions can be entirely satisfactory, and as long as ice formation is not encountered there should be no reason to apply preheat, unless the fueI requires some heat t o aid in vaporization and distribution as indicated by rough running.

CYLINDER HEAD TEMPERATURE

FLIGHT

In attempting to obtain cylinder temperature control by means of mixture regulation, the operator must bear in mind that a decrease in cylinder head temperature is not always obtained by advancing the mixture control from automatic lean to automatic rich. On many engines carburetors are so adjusted that the automatic-lean position provides approximately best economy mixture strength in the cruising range. At corresponding powers the automatic rich position will result in approximate lean best power mixture strength. Under these conditions in the cruising range automatic lean will result in cooler cylinder temperatures than automatic rich. Above approximately 70?& power automatic rich will always result in cooler cylinder temperatures than automatic lean.

LUBRICATION

The specified maximum cylinder head temperature for continuous operation is established to ensure prevention of detonation and to maintain the strength of the head and piston material in spite of the high combustion temperatures resulting from a lean mixture. Exceeding this temperature limit with lean mixtures will result in detonation which will materially weaken the cylinder heads and pistons and reduce the life span of the engine. Engine life will be appreciably increased by maintaining cylinder head temperatures below 400 F (200 C), and every effort consistent with aircraft performance should be made to operate below this temperature.

The procedures outlined under Climb apply equally to Level Flight.

DESCENT Normally, descent can be considered a continuation of level flight, maintaining the same conditions of power, bmep, and temperature control. If constant power output has been maintained by increasing the rpm when the critical altitude

for the normally selected rpm and manifold pressure was passed, maintain the desired power during the descent by reducing the engine speed by 50 rpm for each in. Hg gain in manifold pressure. After passing the normal setting critical altitude, regulate power by throttle adjustment.

PWA. 01. 100


GLIDE AND APPROACH

The transition from cruising descent to glide and approach occurs a t the time that the aircraft is slo the undercarriage and when ente airport flight pattern. From this point until actually landing, the pilot will want the engine in a condition to respond immediately to any demand for power, from a moderate increase for maneuvering to full take-off rating in case

I,

PRESSURE rated rpm may be required r is needed, the propeller advanced to a position result in 85-90% of Normal when the propeller is governe control is not advanced to the osition, a s a sudden throttle result in serious overspeedamount to 600-800 rpm ers. If the need for full , advancing the throttle 11 give immediate power response and he rpm control can be adjusted later. With two-position propellers, the rpm control should be placed in the high rpm position. With either two-position or constant speed propellers, the control position should be adjusted a t an aircraft speed sufficiently low so that overspeeding does not result.

120

- MANIFOLD PRESSURE

\* 7

\I

As a variable pitch (constant speed) propeller will be in fixed low pitch a used, engine power control tirely by the use of the t

GENERAL OPERATING INSTRUCTIONS MIXTURE

- FUEL-AIR RATIO

FLIGHT OPERATION--GLIDE AND APPROACH The remainder of the flight is a t low airspeed and the airflow is needed for ignition system cooling. In case there is a demand for full power, the engine will receive sufficient cooling to prevent a serious temperature rise, and the cowl flaps can be later readjusted after taking care of more urgent items.

LUBRICATION

Rich mixture will be required for the remainder of the flight to obtain slatiafaetory a c d eratian response and t o ensure engine prdection in case of a sudden meed for full power. CHARGE TEMPERATURE AIR TEMPERATURE

-CARBURETOR Where manual oil cooling control is provided, adjust shutters to maintain the desired temperature condition.

CHECK LIST

Under conditions leading to ice formation, maintain lOOF (40C) carburetor entmnce temperature. In the event of a sudden application of fulI power, the earburetor k t control mu& be r d j u e t e d immediately, as the increase in power will r a u l t in a rapid riare in the available carburetor h a t .

CYLINDER HEAD TEMPERATURE

As in preparation for take-off, a landing check-off list is recommended : - Rich (Automatic Rich) 1. Mixture - 85-90% Normal Rated 2. Rpm 3. Fuel Selector

- Cold; o r lOOF (40C) if

6 . Cowl Flaps 6. Oil Cooler Control

- Trail o r 2 in. open - As needed

- --

The cowl flaps, where used, should be awn dightly (trail gaition on e m ins~llations).

On suitable tank*

4. Carburetor Heat

icing conditions a r e present

*Experience has shown that this item cannot be mentioned too of ten.

121

PRArM' & WHITNEY AIRCRAFT --

.---

.. .

PWA. 01. 100

---

LANDING

SHUTDOWN

E ; ~ h cowl e flaps are to be opened as soon as it 1s practical during the landing roll to remove any obstruction to the dissipation of the heat stored up in the engine and to be prepared for taxiing.

With the cowl flaps fully open t h e engine should b e i.dled for a sufficient length of time to ml the cylinder head temperatures t o below limiting shut-down temperatures 400 F (200 C) . Shutting down a warm engine results in leaving en exwmJv@:a w n t i of k t $bred in the mass ef the engine., with no means of conducting i t &way e m p t by mnv@ction currents.

B~uttii-egGE the engine is accomplished as f dlows : 1. Shut a 8 b m t e r pump (if applicabIe). 2. Move mixture control t o idle eut-off or full lean, with the throttle open to any idling speed. N o t e : I t is not necessary to open the thrattle as the engine cats off.

3. When the engine stops rotating, t u r n off the ignition 4. Lmve crawl flaps fully open for a t least 15 minutes.

f

VDER HEAD TEMPERA URE L I M I T

SPARK PLI INTAKE MA PUSH RW MOUNT Pi

'LBOWS 'OLD SEALS ER TUBE SEALS ?TALS

Fig. 41 -Engine Temperatures After Shutdown


LIGHT OPERATION-LANDING TO SHUT DOWN

HEATED AIR

HEATED AIR CONF/NED

BAFFLE SEAL ACTS AS BARRIER TO CIRCULAT/ON

Even though the engine is properly cooled prior to shutdown, i t still contains a large amount of heat, and it takes a considerable time to remove the heat that is stored in the mass of metal. The only means of removal are the convection currents set up within the cowl. Cool air enters the lower portion of the powerplant and rises to the top of the engine compartment. Unless a ready exit is available, this heated air remains static a t a temperature which will damage the ignition harness.

under ground testa.

As the mixture control lever is moved into idle cut-off or full lean, an opportunity is afforded to check the idling mixture as discussed

Opening the throttle, with the ignition cut, leans the mixture to a point where combustion cannot be maintained.

If the engine does not stop with the mixture control in idle cut-off or full lean: 1. Close throttle.

2. Shut off booster pump (if applicable). 3. Cut ignition. 4. Slowly open throttle.

(5. Have idle cut-off checked)

"

As there was

systems must b supercharging ca of t h e engine's operat directly affected by

charger controls. ground by t h e "skin of his teeth" not always bliss. tration the pilot violated the

into expecting a p r with t h e erudition of deal with t h e subject using on1

in power by use of under conditions cance a s t o operation of t h e introduce new factors whi

PURPOSE OF SUPERCHAR

"B" is capable of 235 mph. must be superior t o "B." H

y stage could have

t, he did not know

parison is of little value. At what altitu airplanes "A" and "B" attain these ma be able t o go 285 mph and "B" can do 235

variations in pe altitude on airpl ane spec?d, and in the ability, or lack of ability, of the engrine t o maintain a given . * altitude of operaticIn. The power t o the desirea

.,

ADVANCED SUPERCHARGING

PURPOSE OF SUPERCHARGING

TRUE AIRSPEED

-MIL ES PER HOUR

Fig. 42 -Airplane Performance vs Altitude

degree of supercharging is of greatest importance in enabling the engine to deliver its power a t altitude.

thrust. The variation in airplane forward speed with altitude at various percents of power is shown in Fig. 42.

AIRPLANE PERFORMANCE VARIATION WITH ALTITUDE If, as altitude increases, engine power can be maintained, a gain in level flight high speed will be obtained a t the rate of, roughly, 1% per thousand feet. In each cubic foot of air a t higher levels there are fewer particles of air to impinge against and retard the airplane, whereas the constant speed propeller makes it possible to sustain approximately constant

ENGINE PERFORMANCE VARIATION WITH ALTITUDE As the power delivered by an engine varies with the amount of air it consumes, the performance will fall off a s altitude increases when operating with a fixed throttle position and constant rpm. The volume of air drawn into the cylinders will remain unchanged but the weight of this volume and, consequently, the power generated, will diminish in proportion to the air density as shown in Fig, 43, This is not 100% of power a s defined by the engine rating; i t is 100% of the power obtainable a t sea level with wide open throttle and a t a constant rpm. W i t h supercharged engines, normal rated power (100% rated power) is usually less than 100% of full throttle sea level power. The relation shown is fundamental f o r a l l engines, supercharged or unsupercharged, provided that the impeller is driven a t a fixed ratio to crankshaft speed.

rl

I.

ALTl

TUDC

Fig. 43 -Engine Performance vs Altitude

TRUE AlRSPEED - M I L E S

PER HOUR

fig. 44 -Combined Airplane and Engine Performance vs Altitude

COMBINED AIRPLANE AND ENGINE PERFORMANCE By combining Figures 42 and 43 it is possible to determine the altitude range of performance obtainable with a given airplane and engine combination. For example, point 'ID" on Figure 43 can be directly transferred to Figure 42 by placing i t on a point having the same value of power and altitude. Similarly, the remaining points of Figure 43 are transferred and form --

-

line A-B-C-D-E-F. Line A-F then becomes a curve of the performance obtainable with the airplane and engine combination using the rpm shown and a t full throttle. If the engine speed is the maximum allowable the airplane is capable of the performance shown below line A-F. If performance above line A-F is required, the engine power capacity must be increased. It is the function of the supercharger to enable the engine to have this increased power capacity. 7

100% OF POTENTIAL POWER

D E S I R E D POWER

r n t s POWER L O S S M U S T

BE ACCEPTED

:

2Q

1' POTENTIAL POWER

QI

'4

I I ALTITUDE- FEET

0

01



If the airplane designer's wishes could be followed, the limits imposed by nature would be pushed back and the engine designer would provide a powerplant capable of delivering constant power unvarying with altitude. However, as the facts of atmospheric life are inexorable, the engine performance will invariably follow the path shown in Figure 43. If the desired power must be obtainable at 10,000 feet the engine capable

er must base his design on what power is required a t what altitude. The excess power available a t lower altitudes must be kept in check by proper observance of limits and the operator must accept the diminished output a t higher altitudes.

b. Number of

INCREASING ENGINE POWER CAPACITY As the power developed in the cylinders of the engine depends directly upon the lbs. per hour of air consumed, the need for increased power is engine's air-consuming

Objection-Size:

Extra size of engine results in excessive airplane

L O W A I R CC

h

H I G H A I R CONSUMPTION

Objection-Centrifugal and inertia loads become excessive and valve action is something different than provided for by cam profile.

PRATT & WWITNEY AIRCRAFT

PWA. OI. 100 -

3. - Supercharging

--..

T H I S VOLUME /S THE SAME, AS THIS

lNCHES

CHARGE, VOLUME SWEPT BY PISTON EQUALS THE CHARGE VOLUME

7

'T

d

-3-

/ S DRA Wfi' /NTO THE CYLINDER. T H I S VOLUME NEVER VARIES REGAROLFSS OF ALTITUDE, TEMPERATURE, THROTTLE POS/TlON, OR DEGREE OF SUPERCHARGWG

The power that this charge develops in the cylinder depends upon the volume of atmosphere that is compressed into this fixed charge volume which can be varied by1. Change in charge pressure which can be varied by

a. Throttling Supercharging offers the most efficient means of increasing engine's air-consuming capacity. I t accomplishes the same results as increasing the size or number of cylinders or increasing the rpm but i t accomplishes these results with the least increase in weight or size of the powerplanti llowing table illustrates the variations required size of an unsupercharged engine rder to maintain 1000 hp to various altitudes rpm, compression ratio and

PART T H R O T T L E

ltitude Sea Level 20000Ft. 40000Ft. ltitude Power Capacity 1000 1000 1000 -11 ~ L - - u I . n

U L I L I I I V L L I ~r u w t f r

Capacity at Sea Level 1000

RO

2083

6410

NTRANCE

TO THE PROPELIAER oped in the cylinthe WEIGHT of 'Weight" is emtor that can be Volume cannot be varied.

-FULL THROTTLE

b. Change in carburetor entr sure resulting f


PURPOSE O F SUPERCHAR 2. Change in charge temperature which can be varied bya. Change in atmospheric temperature

SEA

L

HIGH ATMOSPHERIC TEMPERATURE

0,000 FEET

c. Supercharging LOW ATMOSPHERIC TEM PERATURE

JNSUPERCHARGED

n

n NO PREHEAT

,.SUPERCHARGED

n increase in the pressure of t h e charge is a vorabIe factor in increasi Etrge and, consequently, veloped in the cylinders.

WITH PREHEAT

& WHITNEY AIRCRAFT

PWA. 01. 100 cylinder to the propeller and is an important factor in determining the over-all gain from supercharging. POWER DEVELOPED IN T H E C Y L I N D E R S

OWER ABSORBED Y THE IMPELLER

SUPERCHARGED BUT WITHOUT ACCOM PANYlNG TEMPERATURE R l S E A PURELY IMAGINARY CONMTION

POWER DELIVERED TO T H E P R O P E L L E R

---

SUPERCHARGING BASIC FACTORS

ERCHARGED WITH U N A V O I D A B L E O M PANYlNG TEMPERATURE RlSE

charge temperature is an factor in increasing the charge tly, lowers the power proower is developed from the charge nto the cylinder, a flow of power ward the propeller. However, ons must be made from this flow to omer-absorbing agencies within the engine beore the useful balance available a t the propeller haft can be measured. These are: tion horsepower power expended to overcome the internal 'riction within the engine and to drive the various engine accessories. While, in practice, this item is of appreciable magnitude, it does not vary to any great extent because of changes to be considered in this discussion. Therefore, the engine will be ~nnqideredas being frictionless.

The effect on charge pressure and charge temperature by compression, and the amount of power required to drive the impeller are the three principal measures by which a comparison can be made of various superchargers. 1. EfFect on charge pressure Measured by pressure ratio -,the ratio between the pressure at the air entrance to the pressure a t t (manifold pressure PRESSURE RATIO

=

MODERATELY SUPERCHARGED

PRESSURE RATIO=*

30

= 1.5:l

PRES

It is sufficiently accurate

-30 60 --

(



which would not add enough to the accuracy of this presentation to be recognized). For example, a t 19000 feet the entrance pressure is 14.3 in. Hg.

For example: a t 19,000 ft. the temperature of the air a t the entrance is-26'.

MODERATELY SUPERCHARGED MODERATELY SUPERCHARGED

HIGHLY SUPERCHARGED

-

TEMPERATURE RISE +7 O- (-23°)0300C 21 4 PRESSURE R A T I O = ~ = I . S :I

HIGHLY SUPERCHARGED

26.8 PRESSURE RATIO=-= 14.3

TEM PERATURE R ISE = +22O- (-23O)=4S0C

2: l

2. Effect on charge temperature Measured by temperature rise - the difference between the temperature of the charge a t the entrance and the charge temperature a t the supercharger exit. This rise is the inevitabIe result of compression together with unavoidable fluid friction and turbulence.

Besides its effect on power, the charge temperature is a very important measure of the tendency of the charge to detonate. For any grade of fuel a close relationship exists between the temperature and the pressure of the charge which defines the limit of normal safe combustion. If the charge pressure is low it can tolerate a high temperature, or, conversely, if the temperature is low a high pressure can be safely imposed. The limits resulting from this relationship can be shown graphically as follows:

HIGHLY SUPERCHARGED

TEMPERATURE RlSE = 45O-15- =30°c

TEMPERATURE RlSE = 60° -15O = 45OC

As with the case of pressure ratios, i t will be assumed that temperature rises remain constant a t a given rpm regardless of changes in air entrance temperature, airflow as controlled by throttle position or airflow a s determined by the altitude of operation. Actually, there is but a slight deviation from this assumption.

Fig. 46 Charge Temperature vs Detonation

The implication of Figure 46 above is that, a s the charge temperature is increased, it becomes necessary to prohibit the use of that portion of the engine's power capacity that lies in the detonation range.

PWA. 01. 100

NEY AIRCRAFT of operating a t sea level with a carbure-

tion with the ch

h

r L l . W l T REQUIRED I N ABOVE E X A M P L E

In utilizing the supercharger, the engine designer is reaching for the goal of maximum power sustained to the highest possibIe altitude. By supercharging, he is endeavoring to obtain the maximum benefit from the resulting increased- charge density without being penalized excessively by restrictions resulting from the increased charge temperature.

€R L I M I T IS REQUIRED W I T H LOWFR GRdOE F U E L

ALTITUDE

essary Power Restrictions tion-Free Operation

The step by step development of the supercharging systems that follows uses the same engine components tabulated below. All operation is a t constant rpm.

3. Supercharger horsepower

the impeller varies city a t the impeller

I. Power Section Cylinder displacement, compression ratio, spark advance, valve timing and Iift and fuel-air ratio.

pipes. The only variable used is impeller speed with the exception of the two-stage engine.

0:I IMPEL D R I V E Rl

LEVEL

varied to obtain desired a1

ER

based on the premrpm supercharger es with altitude in

,

.

/

POWER REQUIRED / BY T H E I M P E L L E R

POTENTIAL POWER A V A I L A B L E AT THE PROPELLER

POWER PRODUCED I N THE CYLINDERS

WEIGHT OF A I R C O N S U M E D PER HOUR 1 T H E MEASURE OF POWER PRODUCED I N T H E

CHARGER PUM PING CAPlClTY

T H E MEASURE OF

VOLUME OF AIR

n nCHARGE PRESSURE THE MEASURE OF \ PRESSURE RATIO

<

CHARGE TEMPERATURE T H E MEASUREOF TEMPERATURE RISE


PWA. 01. 100 RESULTS OF SUPERCHARGING

TYPICAL UNSUPERCHARGED (NORMALLY ASP I RATED)

Supercharging does not change the fundamental performance characteristic of engine performance with change of altitude. A supercharged engine obtains 100% of its total power capacity at sea level. At any altitude, full throttle power a t constant rpm follows the same relationship, as shown in Fig. 43.

1

w

PRESSURE RATIO 0.9:J TEMPERATURE RISE 0 C

T YPlCAL SUPERCHARGED

This will be understood if it is realized that the supercharger has merely supplemented the fixed volume pumping capacity of the cylinders with an additional fixed volume of capacity. With constant rpm and wide open throttle the supercharged engine will always draw in the same volume of atmosphere. The power will diminish with increasing altitude as the weight of this fixed volume diminishes. The engine designer could have obtained the same results by increasing the size or number of cylinders or, possibly, by increasing the rpm.

PRESSURE RATIO a:? TEMPERATURE R14E 45%

FULL THRBTYLECHARACTERISTICS AT SEA LEVEL

The essential result of supercharging, then, is that the pumping capacity of the engine has been increased and, with it, the capacity of the cylinder to produce power. The supercharger illustrated is a single-stage, single-speed supercharger because 1. The compression is accomplished in one operation or stage 2. The impeller is driven a t one fixed ratio to the crankshaft

ALTITUDE

-

' , ..

FEET

Cornperrisen ef Power

1 I

$ I

s bl

POTENTIAL PERFORMANCE NOT USEABLE BECAUSE OF OPERATING RESTRICTIONS

i=

. I

7

TRUE A/RSP€ED

- A4 P-li

Effect en Airplane Performance

SINGLE STAGE GROUND BOOSTING

ADVANCED SUPERCHARGING TYPICAL SYSTEM WITH 6:l

IMPELLER DRIVE RATIO PRESSURE RATIO I.5:I TEMPERATURE RlSE

30"~

CHARACTERISTICS AT SEA

LEVEL

PRESSURE R A T I O I.5:l TEMPERATURE RlSE

3O0C

CHARACTERISTICS AT 19,000 FT. ALTITUDE

STEPS IN SINGLE-STAGE DEVELOPMENT 1.

Ground Boosting

The power developed by the engine in this example is the maximum usable sea level performance which can be obtained from t h e basic power section. If the impeller drive ratio were lower'.: the engine would not receive sufficient air to develop this output. If the ratio were higher, the increased charge temperature would require that the maximum pressure be limited in order to prevent detonation. The power available within this restriction would be less than the maximum power obtainable with the ground boost supercharger. The example shown was selected t o illustrate t h a t the given ground boost supercharger provides the means of obtaining the highest possible sea level performance. Any supercharger that increases the power capacity of an engine over that obtainable without supercharging and which does not result in a restriction of operation requiring throttling at low altitude is considered a ground boost supercharger. Engines equipped with ground boost supercharging are used on basic training and light transport aircraft. The simplicity of operation esulting from freedom of restrictions recomnend it for this class of airplanes.

Comparison of Power

Effect on Installed Weight TRUE A/RSPEED - M P H

Effect on Airplane Performance

Weight hp @ S.L. Rating wt/hp @ S.L.

Unsupercharged 900 775 1.16

GroundBoosted 1000 1000 1.00

'

PRATT & WHITNEY AIRCRAFT wrln

PWA. 01. 100 2.

r . 1 I M ~ L L L E R DRlVE R A T I O

'

PRESSURE RATIO 2:) RISE 45°C

-

-the

Low Altitude Supercharging

Stepping up the speed of the impellc; provides the engine with the capacity t o consume more Qlcubicfeet of air a t any altitude. Near sea levei engine has the power capacity to exceed the performance that has been set up as a goal by er. However, because this 1degree of :in g results in an increased temperature rise, the maximum usable power, or rating, ufficiently low so that the pressure .ned with the resulting high1er charge temperature will not produce detonation. The maximum usable power within this Iimitation is Iower than the usable power of the ground-boosted engine provided with the optimum impeller drive ratio and using the same grade of fuel. In selecting this degree of supercharging the designer has accepted the small loss in performance close to sea level in order to gain considerable altitude output. The engines using this degree of superchargual!ly sustain their maximu~ n ratings 100 feet. This characteristiLC recom.r installation on transport a.ircraft in r to provide the required take-off performrts located above sea 1e1.el and t o cruising performance a t moderate altitudes. The usual cruising powers of these engines are sustained to 8000-14000 feet which provides a satisfactory altitude range for medium distance airline operation without cabin supercharging. With engines equipped with a lower degree of supercharging, engine speed is the only major restriction in operation. With the introduction ~percharginga new limil; must be .is new limit is charge pre ssure and is indicated in the cockpit by the manifold pressure gage.

Effect on Installed Weight

.:

Weight hp 6 S.L. Rating wt/hp 6 S.L.

Ground Boosted 1000 1000 1.00

* * * * T R W AIRSPL,,

MPH

Pff ect on Airplane Performance

hp @ 4000 ft. wt/hp 8 4000 ft.

875 1.14

Altitude Supercharged 1000


SINGLE STAGE ALTITUDE SUPERCHARGERS

TYPICALSYSTEM WITH 10:l IMPELLER D R I V E R A T I O PRESSURE RATIO 3 : l TEMPERATURE RISE 6dc

3.

High Altitude Supercharging

Further increase in impeller speed still further increases the volume of atmosphere drawn into the engine and the potential power capacity has been correspondingly enlarged. However, the supercharger heat rise is now of such magnitude t h a t t h e maximum power restriction is severe and the performance a t low altitude is considerably reduced.

FULL THROTTLE CHARACTERISTICS AT SEA LEVEL PRESSURE RATIO 3:I

n nTEMPERATURE RISE 6

6 ~

These characteristics recommend the use of engines equipped with this degree of supercharging to those airplanes requiring maximum altitude performance but which are able to take off and climb with relatively low power. Fighters are examples of this class of aircraft. Effect on Installed Weight

715

m

Supercharger Low Altitude High Altitude 1000 1000 Weight hp @ S.L. Rating 975 940 wt hp @ S.L. 1.03 1.06

BHP

* * * * FULLTHROTTLE CHARACTERISTICS AT I 9 0 0 0 FT.

hp 8 12000 ft. w t hp 6 12000 ft.

733

1.37

LIMIT OF IMPELLER SPEED

AL Ti N U L

- FEET

Comparison of Power

TRUE AIRSPt-ED

-M P H


It can be seen t h a t each step of increasing impeller drive ratio has made it necessary to impose more severe restrictions upon the brake horsepower t h a t can be safely used a t low altitude. Eventually, this restriction will become so severe t h a t sufficient power will not be available for satisfactory performance a t sea level. Between the limits of optimum ground boost impeller speed and the limiting maximum speed, the engine designer has the choice of an infinite number of impeller drive ratios. His selection of any one is based on the type of performance which must predominate. If maximum output is required at sea level the consequent lack of performance a t altitude must be accepted. Conversely, if maximum altitude output is essential a reduction in maximum usable power is unavoidable.

PWA. 01. 100


ALTITUDE F E E T

Fig. 48 -Engine Altitude Performance With Different Impeller Drive Ratios

e maximum single-stage altitude per-

The comparative powers obtainable the usable range of impeller gea above.

Both ratios provide the engine with the same stics of air volume consumed, pressure perature rise and supercharger horseower as with .the correguonding single ratio ......::

ADDITIONAL SINGLE-STAGE DEVELOPMENT 'The engine designer cannot reach t h e goal maintaining constant power a t any altitude wi a single-stage supercharger using a single fixedj impeller ratio. Each time he reaches for more' altitude performance he loses ground near sea level. If he is t o realize the possible gains of high impeller ratios without losing the advan-. tages of the lower degree of supercharging he. must devise a drive system which allows the impeller to be driven a t a s p e d more nearly1, optimum to the altitudle of operation. This ne-: cessity has brought about the development of . the multiple and variable speed impeller drives.,

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Comparison of Power

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER By providing a two-speed impeller drive a n 8 selecting ratios from Fig. 48 giving best per-, formance a t two desired altitudes, greater alti-: tude range of airplane performance is obtained. The low ratio is selected to give the desired output near sea level and can be any degree sf supercharging from ground boost t o t h a t for low altitude. The high ratio is usually selected

TRUE AIRSPEED - M P H

Iffect en Airplane Performance

ADVANCEDSUPERCHARGING The single-stage, two-speed supercharger enables the engine to develop high power for takeoff at or near sea level and, a t the same time, permits the engine to maintain high airplane performance throughout the medium altitude range. I t will be noticed that in this case the maximum permissible high ratio power is greater than that of an engine having a single-speed drive of the same ratio. The gain is possible because the use of the higher ratio is reserved for altitudes where it will permit the engine to develop more power than with low ratio. For example, a t 10,000 feet (Point A ) , the perniissible engine power can be based on a charge temperature of -5 C plus the supercharger temperature rise. The same impeller speed a t sea level will produce a charge temperature of 15 C plus supercharger temperature rise. The 20 C more favorable charge temperature permits a greater charge pressure and, consequently, more power. The availability of this additional power makes it imperative that the high ratio be used only where advantageous. Premature engagement of high ratio will very probably result in detonation. This premium of extra power when using high ratio power makes it necessary to assign two different limiting carburetor air temperatures with single-stage, two-speed engines. The higher limit is allowed with low ratio supercharging onIy. When high ratio is engaged, the lower limiting entrance temperature must be observed. It is only by the use of these carburetor air temperature restrictions that this additional high ratio power rating can be offered and i t is essential that these considerations be understood when using carburetor preheat for ice prevention or when flying in air of extremely high temperature.

SINGLE STAGE ALTITUDE SUPERCHARGERS Recause of the increased versatility of performance possible with the single-stage twospeed engine i t is used on a wide variety of aircraft requiring maximum performance from sea level to medium altitude.

Effect on Installed Weight SingleSingleSneed Iineed ~ r & n d - wit; Eigh Booated Ratio Engine Drive

Two-Speed Low High

Weight Hp @ S.L. Rating Wt/hp @ S.L.

1000 1000 1.00

1000 940 1.06

1040 1000 1.04

Hp @ 12000 ft. Wt/hp @ 12000 ft.

668 1.62

940 1.06

-

1040

-

940 1.11

LIMITATIONS OF FIXED RATIO DRIVES Even the multiple speed drive is a compromise. The fixed ratios are selected to give maximum performance a t specific altitudes and possible performance between these points is sacrificed. For example, the engine is provided with an 8 :1 drive ratio and is rated a t 750 hp a t 2500

rpm. The highest altitude to which this performance can be maintained is 10,000 feet which is termed the critical altitude for this power and rpm. At this rpm and a t full throttle the supercharger pressure ratio is 2.25:l. When operating a t 5000 feet the impeller speed has not changed and, hence, its pressure ratio remains the same a s a t 10000 feet. However, as the atmospheric pressure has increased to 24.9 in. Hg, full throttle operation would result in raising the manifold pressure to 56.0 in. Hg which, in combination with the increase in charge temperature to 55 C, would result in detonation.

The addition of the two-speed drive involves little additional weight to the powerplant (2040 Ib) and no additional drag. Operation is somewhat more involved as ratings, limiting powers, rpm, manifold pressures and bmep are not the same for each ratio. In effect, the pilot is operating two different engines and must be aware of the differences in order to maintain proper conditions and to obtain the most efficient utilization of the powerplant. ALTITUDE

- FEET

PWA. 01. 100


In order to restrict the engine output i t is necessary to use the throttle so that the pressure a t the face of the impeller does not exceed that obtained a t this same point when operating a t 10000 feet and with full throttle.

10,000 F E E T F U L L THROTTLE

drive the impeller a t 5000 feet is lower with 7:l than with 8:l drive ratio. This represents a direct gain in power delivered to the propeller. By providing a variable speed impeller drive the engine designer enables the single stage engine to deliver the maximum possible performance a t all altitudes and provides i t with the greatest operating flexibility.

5000 F E E i i

PART THROTTLE

Cempwison d Power

As the impeller speed remains con&m%. power required to drive the impella Its smchanged.

THE SINGLE-STAGE VARIABLE SPEED SUPERCHARGER If, instead of restricting the supercharger output by means of the throttle, the necessary regulation could be obtained by means of varying the impeller drive ratio while holding full open throttle an appreciable decrease in power required to drive the impeller will result. Taking the previous example a t 5000 feet, the power delivered to the propeller can now be increased to 775 hp because the power required to

TRUE AIRSPEED


Effect on Installed Weight

Weight Hp @ S.L. W t / h ~@ S.L. Hp @ 6500 ft. Wt/hp 8 6600 ft.

1040 1000 1.04

- FEET

1-stage Variable Speed 1070

* * * *

940 1.11

* * * *

Hp @ 12000 ft. Wt/hp 8 12000 ft.

ALTITUDE

-M P H

9.40 1.11

The variable speed supercha~gerresults in a performance which is the combination of the maximum altitude outputs obtainable with an infinite number of drive ratios varying from maximum ground boost to limiting maximum. A t any impeller speed the supercharger has the same characteristics of air consumption, pressure ratio, temperature rise and supercharger horsepower as with the corresponding single ratio drive.

TWO STAGE SUPERCHARGERS

ADVANCED SUPERCHARGING The variable speed drive provides the single stage engine with the maximum versatility of operation. The cost in weight is low (60-100 1b) and no increase in drag is involved. As its operation is, by necessity, automatic, control is greatly simplified. These qualities recommend its use on all aircraft requiring maximum performance from sea level to medium altitude.

THE TWO-STAGE SUPERCHARGER A single-stage compressor, even though all components are designed for the highest possible altitude of operation, reaches its limitations in the medium altitude range. These limitations are : 1. Size

While the examples previously shown are discussed on the basis of maintaining constant the physical dimensions of the supercharger while varying only the impeller speed, in actual practice the entire size of the compressor is increased. Entrance passages, impeller diameter and width, diffuser and collector undergo increases in size in order t o handle efficiently the larger volume of air. Accordingly, a supercharger with proper sized components for high altitude will not be capable of efficient performance near sea level. As the volume of air t o be pumped becomes greater and the diameter of the impeller is increased, so does the size of the cases housing the impeller. Engine installation requirements make it essential to keep this engine section below a definite maximum diameter. 2. Temperature rise

As the pressure ratio increases, the resulting temperature rise becomes excessive and imposes impractical limitations on usable brake horsepower.

The foregoing examples t r e the d m n t of the singte-stage supe~hargerf m the mrhtiom in pemforrazance a a l p l a b k by man6 of different types uf i m p l h drive. In actual pmetiee the designer utilizes a e m d d d b n u m k of w i b i e changes in the design of the impeller, induction pasaim end diffuser in ~rdm to oMsin the m w t eflicient combination to give the &sired performame in a specified altitude range. The considerations involved in the design of a low altitude supercharger, which handles emall volumes of air, differ wrnmhat from the design of a supercharger required to make possible maximum engine performance at 15000 f e e t However, them b c t o r s & not haw appreciable effeet upon the method of operatim and the precautions necesarary for protectfon of the engine.

These limitations are avoided by dividing the supercharger into two separate parts and accomplishing the compression in two steps. This system is called the two-stage system as the compression is accomplished in two steps 0.1stages. The stages are connected in series and the limitations of the single-stage system are avoided as follows: 1. Size Two stages of compression are provided. The main stage is identical to the supercharger of the single-stage engine and, on the system used by Pratt and Whitney Aircraft, it is possible to operate near sea level using the main stage only. By this means, the size of the main stage components is optimum for low altitude performance. The auxiliary stage is designed to handle efficiently the larger volume of air a t high

I

PWA. 01. 100

PRATT & WHITNEY AIRCRAFT altitude. It delivers the air to the main stage at approximately sea level pressure so that, a t all times, the main stage is functioning near its point of best perf ormance.

INTERCOOLER

By using two small impellers having the same airflow capacity a s one of large size, the diameter of the supercharger section is held within reasonable installation limitations.

SINGLE LARGE SUPERCHARGER

T W O SMALL SUPERCHARGERS

B O T H PUMP THE SAME QUANTITY OF A I R

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1

MAIN STAGE (PRIh&RY> IMPELLER-

Main Stage Same as single-stage, single-speed. This sMge js always engaged. The impeller drive ratio 1s selected t o give the desired performance near sea level without the use of'the auxiliary stage. For take-off, maximum performance near sea level, and cruising economy, operation is with main stage only.

Fig. 49

Auxiliary Stage Engaged when performance requirement s cannot be met by sole use of the main stage.

- The Two-Stage System with Gear Driven, Two-Speed

Auxiliary Stage



2. Control of temperature rise

creased airplane size, plus the drag involved in passing cooling air through the intercoolers, detracts from the aerodynamic cleanliness of the installation. Taking two similar airplanes, one using singlespeed supercharging, the other two-stage, and both designed to obtain the best utilization of their powerplants, the airplane with the singlestage will be lighter and cleaner and will have

SINGLE LARGE SUPERCHARGER

T W O SMALL SUPERCHARGERS

W I T H O U T COOLER

WITH INTERCOOLER

-

By cooling the charge between the steps of compression the temperature a t the intake port, is held within safe limits. This between-stage or intercooling is utilized on all Pratt & Whitney Aircraft two-stage systems but it is not necessarily used on all two-stage engines.

EFFECT OF TWO-STAGE SYSTEM ON AIRPLANE The addition of the auxiliary stage to the single-stage engine involves penalties in the form of increased engine weight and size, and additional weight, size and drag to the airplane. In addition to the added engine weight, the supercharger must be charged with the weight of the intercoolers, ducts and extra controls. The in-

TRUE AIRSPEED- M P H

superior performance in the altitude range where its engine can maintain equal power.

PWA. 01. 100

PRATT & WHITNEY AIRCRAFT T Y P I C A L SYSTEM W I T H T H E FOLLOW1 N G I M P E L L E R DRIVE SATIOS M A I N STAGE (NEUTRAL) 6:l AUXILIARY STAGE LOW 6 d : I AUXILIARY STAGE HIGH 8:l rlNTERCOOLER

BHP

S E A LEVELM A I N STAGE

INTER COOLE

CHARACTERISTICS AT^ MAIN PLUS AUXmILIARYSTAGE

30,000FT-

A Typical T w o Stage Sydem

TRUE A/t?SPEED

-M PH




TWO STAGE SYSTEMS Gear-Driven Auxiliary Stage a. Fixed Ratiq Drive The use of the two-stage, two-speed supercharger allows the engine t o deliver its maximum output a t or near sea level and a t the same time makes it possible to maintain high output to higher altitudes than could be obtained with the most advanced single-stage system. These characteristics recommend the use of this type of compressor on installations requiring a uniform high performance through a wide range of altitude without unduly penalizing the compactness of the installation. Many of the most effective fighters used in World War I1 utilized engines equipped with this type of supercharger. Efficient operation of the two-stage, two-speed supercharger places greater demands upon the pilot than singlestage systems. In effect, three different engines are selectively available and in each stage, different limitations of power, bmep and carburetor air temperature are called for. Efficiency is largely determined by the knowledge of the correct altitude range in which t o use each degree of supercharging so that the maximum potential performance or economy can be realized.

mum efficiency can only be obtained by a thorough knowledge of the principles applying t o its operation Effect on Installed Weight

.

Wt. Engine Extra Items, Intercoolers, Ducts, etc. Total Power @ S.L. Rating Wt/hp @ S.L. Power @ 30000 Ft. Wt/hp @ 30000

2-Stage, 2-Speed 1320

-

-

250

* * * 1070 *

1570

1000 1.07

1000 1.57

483 2.22

890 1.76

* * * *

b. Variable Speed Drive The two-stage gear-driven supercharger can be provided with a variable speed drive similar t o that used on the single-stage engine. It is possible to use this type of drive on both main and auxiliary stages or to drive the main stage impeller a t a fixed ratio reserving the variable speed operation for the auxiliary supercharger. The advantages realized are the same as with the single-stage supercharger.

0

P

2

TWO SPEED

B 2 P

For example, if the shift from "neutral" to "low" is made prematurely, considerabIe power is lost as "neutral" will furnish more power to the propeller a s long as i t can maintain the manifold pressure to which '"ow" is limited. If the shift is deIayed by allowing the full throttle manifold pressure to fall off below the proper shift value while in "neutral," possible performance is again lost by not taking advantage a t the right time of the increased permissible power capacity of "low." The same considerations apply when shifting from "low" to "high." While rough "ruIe of thumb" procedures for shifting have sufficed to obtain a reasonably effective use of this equipment, maxi-

1-Stage, Variable 1070

'a

L ALTITUDE- FEET

Comparison of Power

I.

4

TWO-STAGE

9

I.

2

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TRUE AIRSPEED

-M P H

Efiect on Airplane Performance

While the mechanism required t o drive and control the two-stage, variablespeed supercharger represents in-

PRATT &

WHITNEY AIRCRAFT

creased complications, i t results in relieving the pilot of considerable responsibilities for decisions regarding the use of the various compressor stages. Operation is entirely automatic and calls for little more cockpit attention than would be required by a singlestage engine. c. Other Arrangements of Gear-Driven Two-Stage Superchargers The two-stage arrangement shown in Fig. 49 is the one that has been used successfully on Pratt and VVhitney Aircraft engines. Other engine manufacturers have used other arrangements success full^.^ The selection of the system depends upon the engine configuration and the installation problems involved.

PWA. 01. 100 2. The Exhaust Turbine-Driven Auxiliary Stage

The exhaust turbine-driven supercharger is basically the same as the auxiliary stage of the gear-driven two-stage supercharger. The impeller, diffuser and collector are designed to accomplish the same function as a gear-driven auxiliary stage. The essential difference is that with the turbo supercharger the impeller is driven by a turbine utiIizing the energy available in the exhaust gases. The relation between the auxiliary and main stage, intercoolers and carburetors are the same in both systems. A turbine-driven supercharger is, then, a variable speed auxiliary stage compressor and its use allows the maintenance of sea level power to high altitude within the limits of allowable turbine speed and over-all temperature rise.

SUMMARY

B O T H STAGES ALWAYS t N G A G E D

AUXILIARY S T A ~ E ENGAGEMENT SELECTIVE

The abcwe comloinations can be rearranged almosr, indefinitely with the carburetor before the auxiliary stage, interstage or after the main stage. The nterstage or after the 1

.

Supercharging offers the means of obtaining wide varieties of application of a given basic engine power section to different altitude requirements. The engine designer, utilizing the same cylinders, pistons, main cases, etc., is able, by the addition of the proper degree of supercharging, to make the same basic engine fit the peculiar requirements of several types of aircraft. To accomplish the same results by means of unsupercharged engines, the engine designer would have to provide a completely different engine for each aItitude need. The advantages of manufacturing economy and in supply and maintenance are obvious. Each additional increase, in the degree of supercharging, enlarges the restrictions necessary in order to insure safe engine operation. It is essential that the operator understand the necessity for these restrictions and base his operating procedures on a well-founded knowledge of the effect of the supercharger used on engine operating efficiency and safety. For the users of engines e of the several superchargers, Aircraft has publications aval

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A I R C R A F T

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EAST HARTFORD 8. CONNECTICUT, ZT. S . A .

THE SIMQLbSTAGE, KWQdRElED SUPIEWCMARGER AND ITS QIPEWATIQN

DECEMBER, 1945

INSTALLATION ENGINEERING

W H I T N E Y

P R A T T

A I R C R A F T

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TABLE OF CONTENTS

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER OPERATION OF THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER . . . . . . . GROUND OPERATION . . . . Engine Ground Operation Selector Valve and Clutch Check

1

. . . .

4

. . . . . . . .

4

FLIGHT OPERATION . . . . . Take-Off Climb Cruising Descent and Dive Glide and Approach for Landing Carburetor Air Temperature Limits HOW TO SHIFT . . . . To Increase Impeller Ratio To Decrease Impeller Ratio

.

. . . . . . .

5

. . . . . . . . . .

6

. OPERATIONAL CHARACTERISTICS Sludge Formation Manifold Pressure Surge and Pulsation

6

. DETERMINATION OF SHIFT ALTITUDES Determination by Operating Curves Effect of Ram Results of Shifting a t Wrong Altitude Determination by Pilot's Check Chart or Power Plant Chart Determination by Rule of Thumb

7

NOTICE : The following pages contain a general description and a discussion of the operating characteristics of Pratt & Whitney Aircraft singlestage, two-speed supercharged engines. Specific operating procedures for engines installed in military aircraft are contained in Pilot's Handbooks. The information contained herein should not be construed as superseding similar data in applicable Pilot's Handbook.

THE SINGLE-STAGE, T W O - S P E E D SUPERCHARGER A N D I T S O P E R A T I O N

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER The Single-Stage, Two-Speed System.-High performance is required of many aircraft over a wider range of altitudes than is possible with the single-stage, single-speed supercharging system just described. To meet these demands, it has been necessary to devise larger and more complex systems. One of the simplest of these is the single-stage, two-speed system. It differs from MANIFOLD PRESSURE GAUGE

INTAKE PIPE-

COLLECTOR-

&

CARBLJRETOR AIR TEMPERATURE GAUGE

LOW RATIO

. .

incorporation of a gear shifting mechanism by means of which the single impeller may be driven a t either one of two different ratios to crankshaft

speed: low or high. Its basic features a r e illustrated schematically in Fig. 1 The gear shifting mechanism consists of two oil-operated, cone-type clutches and a selector valve. One clutch engages the low impeller ratio drive gear, the other the high ratio gear, while the manually operated selector valve directs engin.e pressure-oil to engage either one clutch o r the other: he complete mechanism is sketched, and its operation briefly described in Figs. 3 and 4 on pages 2 and 3. Typical low and high impeller gear ratios range from 7.15 :1 ( L ) and 8.47 :1 ( H ) to 7.63 :1 ( L ) and 9.89 :1 ( H ) . This arrangement adds little to the weight or complexity of the single-stage, single-speed system, and does not increase either frontal area o r drag. At the same time it makes possible high performance both a t sea level and at medium altitudes. At or near sea level the engine is operated in the low impeller gear ratio, thus keeping to a minimum the temperature rise through the supercharger and the power lost to the impeller. At an altitude determined by the conditions under which the pilot is operating the shift is made from the low to the high gear ratio. The increased airflow and higher manifold pressures available in the latter ratio make possible continued high performance to still greater altitudes, and this gain is realized without any appreciable sacrifice of sea level power. The performance characteristics of a typical single-stage, twospeed engine operated a t the maximum permissible bhp are illustrated in Fig. 2 below. POWER LOST TO FRICTION AND ACCESSORIES POWER ABSORBED IN DR/YING SUPFRCHA RGEU

POWER GAINED BY

ALTITUDE Fig. 2

- Single-Stage,

Two-Speed Performance

1

a @

WHEN IN LOW /MPELLER GEAR RATIO: COCKP/T CONTROL IS F/XED TO HOLD SELECTOR V A L K /N POS/TION TO ADM/T ENGINE PRESSURE O/L /NTO LOW MT/O CLUTCH THROUGH PASSAGE /N CENTER OF CLUTCH SHAf7; PRESSURE OIL FORCES LOW RATIO CONE /N ENGAGEMENT WITH FRICTION SEGMENTS C CLU TCH FACINGS ) WH/CH /N TURN ENGAGE LOW RATIO GEAR. PDWER IS TRANSUITTED FROM IMPELLER lNTfRMED/ATE DRIVE PINION THROUGH INTEGRAL CLUTCH SHAFT AND SPLINES TO LOW RATIO CONE AND THENCE THROUGH FRICTION SEGMENTS TO LOW RATIO DRIVE GEAR AND SO JO IMPELLER SHAFT LOW RATIO GEAR.

ENGINE PRESSURE

@

AS LOW RATIO CONE /S FORCED AGA/NST FR/CT/ON SECMENTS. SPRING MOVES HIGH RATIO CONE AWAY FROM HIGH RA T/O FR/CT/ON SEGMENTS. H/GH RATIO DRIVE GEAR RECE/VES NO POWER FROM CLUTCH 9LAf 7; BUT SINCE IT IS ENGAGED w / r H IMPELLEA SHAFT HIGH RATIO GEAR, IT /DLES ON CLUTCH SHAF Z

@

PRESSURE OIL HOLDS FLOATING VALVE AT REAR OF SHAFT THUS BLOCKING O/L PASSAGES LEADING TO CLU K H .

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DRIVE SHAFT

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RAT/O CONES SL/DE ALONG THESE SPLINES. 1

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IMPELLER SHAFT LOW M T l O GEAF

IMPELLER SHAFT HIGH R A ~ OGEAR

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:.,? /--IMPELLER

COPENS AND PASSAGES TO RATIO CONES.

INTERMEDIATE DRIVE PIN CLUTCH SHAFT J

CSEPARATES AND PREVENTS SIUUL TANEOUS ENGAGEMENT OF 80 7H CLUTCHES.)

Y SHAFT

LOW RATlO CREEPER GEAR (ASSISTS DESLUDGING OF CLUTCH1

LOW RATIO DRIVE GEAR

FRICTION SEGMENTS (CLUTCH FACINGS HlGH RATIO CREEPER GEAR C ASSISTS OESLU S I N G OF CLUTCH 1

LOW RATIO CONE HlGH RATIO DRIVE GEAR

k -

DRIVING MEMBERS

PRESSURE OIL

DRIVEN MEMBERS

DRAIN OIL

Fig. 3 - Clutches and Selector Valve

- Low

Ratio Position

-

m

HIGH

WHEN COCKPIT CONTROL IS MOVED TO "HIGH" POSITION : SELECTOR VALVE CUTS OFF SUPPLY OF ENGINE PRESSURE OIL TO LOW RATIO CLUTCH AND DIRECTS OIL THROUW PASSAGES TO REAR OF CLUTCH SHAFT. FLOATING VALVE IS FORCED FOAWARD THUS ADMITTING OIL TO H/GH RATIO CLUTCH AND BLOCKING PASSAGES TO LOW RATIO CLUTCH. HIGH RATIO CONE IS FORCED AGAINST FRICTION SEOWEm ENGAGING HIGH RATIO DRIVE E A R AND BY MEANS OF SPRING FORCING LOW RATIO CONE AWAY FROM LOW RATIO FRICTION SEGMENTS. OIL /N LOW RATIO CLUTCH AND COMMUNICAT/NG PASSAGES DRAIM OUT THROUGH CREEPER GEAR BLEEDS AND SELECTOR VALVE. ENGINE PRESSURE OILII

DRIVING DNVEN -

I

MEMBERS

PRESSURE OIL

MEMBERS

DRAIN OIL -

Fig. 4 - Clutches and Selector Valve - High Ratio Position

-

=

.TION OF '0-SPEE'DSUPERCHARGER low to the high impeller ratio. 4. Watch for changes in oil pressure, manifold pressure, and rpm. Proper selector valve and clutch operation when shifting from "LOW" to "HIGH" is indicated by : a ) A momentary drop in oil pressure, as the oil fills the high ratio clutch (es) b) A rise in manifold pressure of approximately 1in. hg., as the pumping capacity of the impeller is increased. c) An increase in .rpm due to the higher manifold pressure. 5. Open the throttle until the tachometer reads 2000 rpm. 6. Move the supercharger control lever from is, shift "HIGH" back to "LOW"-that from the high back to the low impeller ratio. 7. Watch for changes in oil pressure, manifold pressure, and rpm. Prope'r selector valve and clutch operation when shifting from "HIGH" to "LOW" is indiated by : a ) i A momentary drop in oil pressure, as : the oil fills the low ratio clutch (es). b) A slight drop in manifold pressure, as the pumping capacity of the impeller :' is decreased. c) A decrease in rpm due to the lower manifold pressure. ,. ',. Should erratic changes in oil pressure, manifold pressure, or rpm occur, indicating improper selector valve or clutch operation, repeat the cycle of shifting after first idling the engine a t 1000 rprn for two minutes to permit the clutches to cool. At least two minutes must elapse between each shift cycle (low to high and return) to allow the heat generated in the clutches to dissipate. In making any clutch shift, the control lever must be moved quickly and without hesitation between positions. If the clutches are allowed to slip or drag during a shift, they are likely to be damaged by the heat resulting from continued friction between their surfaces.

.

I

Fig. 5 - Cockpit Quadrant

2. With the propeller governor set in the high rpm (low pitch) position, open the throttle until the tachometer reads 1700-1800 rpm. (With the oil-in temperature a t 40°C and the rpm a t 1700, sufficient oil pressure to operate the clutches is ensured.) 3. Move the supercharger control lever from "LOW" to "HIGH"-that is, shift from the

. '

.

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FLIGHT OPERATION

i

Take-Off.-All take-offs, regardless of altitude, should be made in the low impeller ratio, whether operating normally or with water injection (if the latter is permitted), or during weather which might lead to carburetor icing. The use of the high impeller ratio a t or near sea level will reduce the horsepower available to the propeller a t any given rpm.and manifold pressure, because of the greater power absorbed by the impeller, and will increase the tendency of the charge to detonate, because of the greater temperature rise through the supercharger. Since the impeller lies between the carburetor and the intake pipes, no preheating is accomplished by using the high impeller ratio, and the danger of carburetor ice is not lessened thereby. (An exception to the foregoing is possible in the case of take-offs from high altitude fields, where experience may indicate the advisability of taking off in the high impeller ratio to obtain maximum performance, provided free air temperatures on the ground are sufficiently low to ensure freedom from detonation.) Climb.-Climbs may be made a t any power up to and including Combat Power (also known a s War Emergency Power), provided the shift from the low to the high ratio is made a t the proper altitude, as determined from the operating curves or the pilot's check chart-as explained below-for the particular type of operation selected. If the need for high performance at any given altitude is not anticipated, the use of the low impeller ratio may be continued so long as sufficient power is still obtainable. The penalties for failure to shift a t the correct altitude are discussed below. (See pg. 8 .)

i

Cruising.-The impeller ratio for cruising is selected with reference to altitude and to the type of operation (i.e. percentage of power) desired o r anticipated a t the given altitude. For maximum fuel economy it is generally desirable to operate in the low impeller ratio wherever possible. However, there is actually little difference in cruising performance between full throttle operation in the low ratio a t high rpm and full throttle operation in the high ratio a t low rpm, provided the bhp is the same in both instances.

P W A . 01.6 5

In general, a half closed throttle in the high ratio indicates the desirability of shifting to the low ratio. Descent and Dive.-During cruising descent or normal "letting down" from altitudes a t which it has been necessary to use the high impeller ratio, the supercharger control lever should be shifted to "LOW" as soon as practicable and regardless of altitude-unless descent is limited to the range of high ratio operation, in which case, naturally, no shift is made. On the other hand, if the need for maximum performance is anticipated before descending to the range of low ratio operation, the shift from "HIGH" to "LOW" should not be made until the shift altitude for maximum permissible bhp in the low ratio has been reached. (See discussion of results of shifting a t wrong altitude, pg. 8 .) In the case of very long and gradual descents, economical and efficient operation may dictate the use of a special shift schedule. The low impeller ratio should be used for all dives, except those incident to military tactics a t high altitudes, to prevent excessive wear of impeller shaft bearings and driving gears because of possible engine overspeed conditions. Glide and Approach for Landing.-Glide and approach for landing should be made with the control lever in the "LOW" position. In the event of an emergency, the pilot will then have maximum power available. (An exception to the above is possible in the case of approach to high altitude fields, where experience may indicate the advisability of approaching in the high impeller ratio to obtain maximum emergency performance, provided carburetor air temperatures are sufficiently low to ensure freedom from detonation. ) Carburetor Air Temperature Limits.-Because the heat rise imparted to the charge by the supercharger is normally greater in the high impeller ratio than in the low, the carburetor air temperature limits are lower for the high ratio than for the low. The operator must be careful to observe these limits, otherwise the temperature of the charge a t the intake ports may become so high as to result in detonation. As a general rule, all take-

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER A N D

ITS

OPERATION

F oil by centrifugal action. Accumulating in the clutches it makes shifting difficult and, in extreme cases, impossible. Many clutches incorporate creeper gears that aid in the prevention of this accumulation by intermittently bleeding the pressure oil from the clutches, thus carrying off the sludge formations which tend to cause sticking. Under ideal conditions i t should not be necessary to shift such clutches to keep them in proper working order. However, the formation and accumulation of sludge will vary with the type of operation and the grade of lubricating oil. In particular, prolonged operation in one impeller ratio may result in both clutches accumulating excessive amounts of sludge. As a consequence it may be necessary and desirable to exercise all types of clutches at intervals during flight to ensure their proper operation.

off and flight operations should be performed in the low impeller ratio whenever carburetor air temperatures are above 16°C (60°F). The principal exception to this rule arises in connection with icing conditions. Whenever preheat is applied, carburetor air temperatures may be obtained which are in excess of the limiting values for normal operation. Accordingly, the application of carburetor preheat should be restricted a s far as possible to low impeller speeds in either ratio.

HOW TO SHIFT To Increase Impeller Ratio.-When the proper shift altitude has been reached during climb, the shift from the low to the high impeller ratio is made as follows :

On extended flights, clutch shifts should initially be made every two hours. At each of these times two or more complete shift cycles (i.e. "LOW" to "HIGH", or vice versa, and return) should be made, and the clutch into which the shift has been made should be allowed to operate for about two minutes. At least two minutes should elapse between each shift cycle to allow the heat generated in the clutches to dissipate. (If the exercising shifts are made during operation in the low impeller ratio, and rpm, manifold pressure, and carburetor air temperature are sufficiently low, it will normally be unnecessary to climb to the shift altitude, in view of the fact that operation in the high ratio is of such limited duration.) As favorable experience indicates the absence of sludge formation, the interval between exercising clutch shifts made in flight may be extended.

1. Close the throttle partially to reduce the manifold pressure 3 or 4 in. hg. (Do not change rpm.) 2. Move the control lever from " L O W to

"HIGH" without hesitation. 3. Adjust the throttle to obtain the desired

manifold pressure, as shown by the operating curve or the pilot's check chart. The partial closing of the throttle (Step 1) that immediately precedes the actual shift is to prevent an excessive manifold pressure rise after the impeller has been engaged in the high gear ratio. A few trials should familiarize the pilot with the necessary throttle movement. To Decrease Impeller Ratio.-Shifts may be made from the high to the low ratio without any strain on the clutches, and it is unnecessary to change the throttle setting immediately before the actual shift, since there is no danger of any manifold pressure rise. (Do not change rpm.) Conditions incident to the descent will determine if, and when, the throttle is to be reset following the shift.

Where the operating schedule calls for no extended flights, the clutch checks made in the course of the pre-flight ground test will serve to keep the clutches free of sludge accumulations.

OPERATIONAL CHARACTERISTICS Sludge Formation.-The high rotational speed of the clutches tends to removc sludge from the

fr

a

Manifold Pressure Surge and Pulsation.-The single-stage, two-speed engine is free of troublesome manifold pressure surges during shifting, and pulsation of the type sometimes encountered with two-stage installations is impossible.

f.

P W A . 01. 6 5

PRATT & WHITNEY AIRCRAFT LOW IMPELLER RATIO X

COMPOS/TE LOW RATIO AND HIGH RAT/O ALnTUDE CAL/BRAT/ON

HIGH IMPELLER I000 LIMITING MAN. PRESS.

[Iz

900

LIMITING MAN. -LOW RAT1

$

0

a W

U3 8 0 0 u:

0

I

W

Y

Q

[Iz

m

SHIFT MAN. PR -LOW RATIO

700

600 CRITICAL ALTITUDE -LOW RATIO 2 2 0 0 RPM 8 32 IN.

SHIFT ALTITUDE

-

-

-

500

L.,"""

ALTITUDE

Fig. 6 -Determination f.

CRl TICAL ALnTUDE HIGH RATIO 2200 RPM 8 39 IN HG

-

FEET

of Shift Manifold Pressure and Altitude from Operating Curve

DETERlMINATION OF SHIFT ALTITUDES The altitudes a t which clutch shifts should be made may be determined from : 1. the engine operating or power curves; 2. the pliot's check chart or power plant chart ; or 3. less accurately, by a "rule of thumb". Aircraft operation is a complex function of many variable factors, and it is difficult, if not impossible, to relate shift altitude to any one of them-for example: to rpm, bhp, bmep, manifold pressure, or air speed-by a simple formula. The shift altitude for any particular engine and airplane installation has no fixed value, but will vary with the type of operation desired or anticipated. The term "shift altitude" is itself somewhat misleading, for, as will be discussed, the pilot in practice is generally more concerned with a shift manifold pressure than with a specific altitude in feet. Determination by Operating Curves.-For those familiar with the use of operating or power curves, this method offers the most general solution to the problem of determining proper shift altitudes, and yields the most accurate results. A typical example of its application is given below :

1. The pilot selects a combination of rpm and manifold pressure for low impeller ratio operation which will satisfy the power requirements and be suitable to the type of operation desired. (Assume the combination so selected is : 2200 rpm and 32 in. hg.)

2. From the low impeller gear ratio altitude calibration of the appropriate operating curve the pilot constructs the full throttle, constant rpm line (1) showing the relation between altitude, full throttle power, and manifold pressure in the lour ratio a t the rpm selected (2200 rpm).

Keeping the rpm unchanged ( a t 2200 rpm), the pilot selects some manifold pressure a t which to continue operation in the high impeller gear ratio. The manifold pressure chosen must not exceed .the limits indicated on the high impeller gear ratio operating curve; otherwise its selection is not restricted. In practice it will generally be selected with a view to continuing in high ratio the type of operation followed in low. (Assume the manifold pressure selected in this instance for high ratio operation is 33 in. hg.-1 in. above that used in the low ratio.)

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER A N D I T S O P E R A T I O N 4. From the high impeller gear ratio altitude calibration of the operating curve the pilot constructs the part throttle, constant rpm, and constant manifold pressure line (2) for the particular combination of rpm. and manifold pressure (2200 rpm-33 in, hg.) selected for high ratio operation.

5. The intersection (C) of the full throttle, low ratio line (1) with the part throttle, high ratio line (2) determines : ( a ) the manifold pressure (29.5 in. hg.) in low ratio operation a t which the shift should be made ; (b) the altitude (approx. 10900 ft.) a t which the shift should be made under standard altitude atmospheric conditions, provided manifold pressure is unaffected by ram. This computation is illustrated in Fig. 6 , which is a composite low ratio-high ratio altitude calibration taken from a pair of typical operating curves. In practice only lines (1) and (2), together with the manifold pressures along line ( I ) , are necessary to determine the shift manifold pressure and altitude for the engine for which the curves are drawn. The other lines are drawn merely for the sake of completeness and to indicate the whole cycle of operations, which is as follows : The pilot begins his climb (at A) in the low ratio, a t part throttle, and a t the selected combination of rprn and manifold pressure (2200 rpm and 32 in. hg.) .As he ascends he will gradually open the throttle to keep the manifold pressure constant (at 32 in. hg.) until he reaches the full throttle, or critical, altitude (B) for the low ratio. From here he will continue to climb a t full throttle, with gradually decreasing manifold pressure and bhp. When the low ratio manifold pressure has fallen to the value (29.5 in. hg.) determined by the intersection (C) of the full throttle, low ratio line (1) with the part throttle, high ratio line (2), the pilot will shift from the low to the high impeller ratio, without changing the rpm, and readjust the manifold pressure to the value (33 in. hg.) selected for high ratio operation.

Ascent will be continued in the high ratio, a t part throttle, and a t the selected combination of rprn and manifold pressure (2200 rprn and 33 in. hg.), with the throttle being gradually opened, until the critical altitude for the high ratio (D) is reached. Above this, operation will be a t full throttle with decreasing manifold pressure and bhp. Effect of Ram.-In performing the foregoing computation the possible effect of ram was neglected. In practice, the manifold pressure a t any given altitude, in either full or part throttle operation, will generally be higher than that shown on the operating curve, because of the velocity of the air entering the carburetor air intake scoop. The amount of ram is not constant, but depends on airspeed and the type of installation-in a well designed installation a particular manifold pressure may be obtained as much as several thousand feet above the altitude shown on the power curve. In actual operation, therefore, where the precise effect of ram on manifold pressure and altitude is not known, the pilot will find it more practical to relate his shifts to manifold pressure rather than to altitude. I n other words, the pilot will continue his climb ( a t full throttle) above the critical altitude for the low ratio until the low ratio manifold pressure has fallen to the value determined from the operating curve (as explained in the foregoing section). At this point he will make the shift, and readjust the manifold pressure to the value previously selected for high ratio operation. The change of impeller ratios is thus made with reference to the manifold pressure gauge rather than to the altimeter, and, consequently, is independent of the effect of ram. Results of Shifting at Wrong Altitude.-If the shift is made below the proper altitude, the result is a loss of available power. For purposes of illustration, assume the manifold pressures selected for low and high ratio operation are approximately equal-the rprn being the same in both ratios. As pointed out in the discussion of supercharger limitations : at any given combination of rpm and manifold pressure, more power will be absorbed by the impeller as its speed is increased. Accordingly, at any altitude below the shift altitude less bhp will be delivered to the pro-

.

m-.PRATT

&

1000

-

WHITNEY AIRCRAFT'

-

.

.

PWA. 01. 6 5

'

POWER LOST BECAUSE OF EARLY SHIFT

POWER AVAILABLE BUT NOT USED BECAUSE OF LATE SHIFT

.. SHIFT MANIFOLD PRESSURE

a

0 I

a

m 800

EARLY SHIFT

-LATf

5000

lq000

ALTITUDE

SHIFT

14000

-

2qOOO

24000

FEET

Fig. 7 - Results of Shifting too Early or too Late

peiler shaft in high ratio operation than in low ratio operation, as shown in Fig. 7 . A further objection to shifting below the proper altitude lies in the greater temperature rise imparted to the charge by the impeller in the high gear ratio. !At low rpm's this may not prove serious, but a t .high rpm's i t may lead to detonation. If the shift is made above the proper altitude :no advantage is taken of the higher manifold !pressure available in the high ratio, and the full , :bhp obtainable is not used, as illustrated above. These considerations apply primarily to shifts :made during climb. During normal cruising de- .. scent power is not a factor of prime importance. !-Accordingly, as noted in the section on flight ??operation, the shift from the high to the low '$ratio may be made a t any point above (but not i :below) the shift altitude. -

!Determination hy Pilot's Check Chart or Power -,:PlantChart.-A number of different pilot's check :charts are available which list, for a series of

4

particular type of operation. So long as a pilot. chart will furnish what is probably the: .+,..=, ,*tn..the ' latter ,', . ,"<.,K

most practical and convenient method of determining proper shift manifold pressures and altitudes. This may be illustrated by the following data taken from a typical check chart. (See page 10.) For example: a pilot wishes to operate a t 60% of Normal Rated power, and for reasons of economy selects the lowest permissible rpm of 1750. He will begin his climb in the low ratio at part throttle and a manifold pressure of 31.5 in. hg. As he ascends he will gradually open the throttle, to keep the manifold pressure constant a t 31.5 in. hg., until he reaches the low ratio full throttle, or critical, altitude (6000 feet, without ram), at which point the bhp will be approximately 660. From here he will continue to climb a t full throttle, with gradually decreasing manifold pressure-. and bhp. When the low ratio manifold pressure has fallen to 26.5 in. hg., the pilot will shift from the low to the high impeller ratio, and readjust the manifold pressure to 28.5 in. hg. If the maaifold pressure is unaffected by ram, the shift alti-tude will be at 10500 feet; but, as previously discussed, the pilot should make the shift with ref-: erence to manifold pressure rather than to alti-

-

I

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER A N D ITS O P E R A T I O N

- -----

--

PILOT'S CHECK CHART"

z'Er I

Power Max. Auto Lean 60

55

50

Press.

RPM 2230 2150 2000 2150 1900 1750 2050 1800 1600 1800 1600 1450

Ylt:

-

Crit.

Alt. without Ram 14500 735 13000 -- 9500 16000 660 10000 6000 16000 10000 605 4000 12500 550 7000 2000

Crit: Alt. BHP

26.0 27.0 29.5

24.5 28.0 31.5 23.5 27.5 32.0 25.0 29.0 33.0

-

SHIFT

. -

% of N. R.

- -- HIGH RATIO --Man. Crit.

---.

Low RATIO

Alt. without Ram

Low Ratio Man. Press.

Press. up to Crit. Alt.

Alt. without Ram

Crit. Alt. BHP

.....

...

.. .

.. .

.....

16000 21500

24.5 26.5

27.5 29.5

700

21500 17500

20000 14500 10500

20.5 23.5 26.5

23.5 26.5 28.5

600

25500 19000 14500

17500 14000 8500

22.0 23.5 27.5

22.5 25.5 29.0

550

255%18500 12000

16500 11500 6500

21.5 24.5 28.5

23.0 26.0 29.5

500

21500 15000 9500

-

1

1

---

*Fuel consumption figures and other data omitted.

6 Table I

- Typical

ratio at part throttle and a manifold pressure of 28.5 in. hg., so long as operation at 60% of Normal Rated power is desired, until the high ratio critical altitude has been reached. Above this, operation will be a t full throttle with decreasing manifold pressure and bhp. (Because of power losses to the impeller, it will be noted that, for any given percentage of Normal Rated power, the bhp available in the high ratio is less than in the low.) If in continuoins level flight the pilot has occasion to alter his type of operation, this may sometimes require a shift in impeller ratios, even though the altitude remains unchanged. For example, neglecting the effect of ram : Assume the pilot in the foregoing illustration is flying at 11,000 feet, a t 60% of Normal Rated power, and 1750 rpm. As indicated on the check chart, this calls for operation in the high ratio at part throttle and 28.5 in. hg. manifold pressure. The bhp

Pilot's Check Chart (Part)

is somewhat below the critical altitude value of 600. The pilot presently finds that a reduction in power is possible, and he decides to operate a t 50% of normal rated power and 1600 rpm. Since the shift altitude for this new combination of power and rprn is 11500 feet, and the pilot is flying at only 11000 feet, he must shift from the high to the low ratio-first reducing his rprn from 1750 to 1600, then making the shift, and finally (since the aircraft is above the low ratio critical altitude, 7000 ft., for the rprn percentage of power selected) fully opening the throttle to bring the manifold pressure as close as possible to the desired figure of 29 in. hg. indicated in the chart. (The order of the actual shifting procedure will depend on the type of operational change and the direction of the shift.) The new bhp will be considerably below 550, as the engine is operating well above low ratio critical altitude for 1600 rprn and 29 in. hg. manifold pressure.

r

PWA. 0 1 . 65

PRATT & WHITNEY AIRCRAFT f--

POWER PLANT CHART* ("Fuel consumption, temperature and other data omitted.) --

Combat Power

Military Power 2800 Manif. Press.

SuperCharger

F.T.

High 4

--

(Omitted)

F.T. F.T. 44.0 44.0 F.T. F.T. F.T. 51.0 51.0 51.0

Normal Rated

Maximum Cruise

RPM

2600

2250

-

I

+ High I

Pressure Altitude

Manif. Press.

SuperCharger

Manif. Press.

SuperCharger

22,000 20,000 18,000 16,000 14,000 12,000

F.T. F.T. 40.0 40.0 40.0

High 4

F.T. F.T. 32.0 F.T. F.T.

High

10,000 8,000 6,000

Low

4 I

I

I

I

t

I

.i

Low

Operating Condition

4,000 2,000 Sea ~ e v e l

1

--

I

II I

+

High

F.T. 42.0 42.0

Low 4

42.0 42.0

I

I I I II I

32.5 32.5 32.5 32.5 32.5

4

+-

High Low

4 - 1

I

I I I

I I

I I

I

i

Low

SPECIAL NOTES Values given are for a particular aircraft and engine model, and are corrected for ram in level flight. Maximum performance shift altitudes for Military, Normal Rated, and Maximum Cruise powers occur where the manifold pressure drops to 41.0 in., 37.5 in., and 29.5 in., respectively, a t full throttle in the. low impeller gear ratio.

T a b l e II

- Typical

The Army and Navy make use of another form of chart known as a power plant chart, a typical example of which is shown in part above. For example : A pilot wishes to operate a t Normal Rated power and 2600 rpm. He will begin his climb in the low ratio ("LOW") a t part throttle and a manifold pressure of 42.0 in. hg. As he ascends he will gradually open the throttle, to keep the manifold pressure constant a t 42.0 in., until he reaches the low ratio full throttle, or critical, altitude-between 8,000 and 10,000 ft. for this combination of rpm and manifold pressure. From there he will continue to climb a t full throttle, with gradually decreasing manifold pressure. When the manifold pressure in the low ratio

Power Plant Chart ( P a r t )

("LOW") has dropped to 37.5 in., as indicated in the second special note, the pilot will shift from "LOW" to "HIGH" and readjust the manifold pressure to 40.0 in. Inasmuch as the pressure altitudes are corrected for ram in level flight, the shift altitude of 10,000-12,000 ft. is a true shift altitude only if the indicated air speed in climb closely approximates that in level flight. As this will seldom be the case, the pilot should make his shift with reference to the manifold pressure gauge rather than to the altimeter, since the manifold pressure will take into account the effect of ram. Above the shift altitude ascent will be continued in the high ratio at part throttle and a manifold pressure of 40.0 in., with the

THE SINGLE-STAGE, TWO-SPEED SUPERCHARGER A N D I T S O P E R A T I O N i

throttle being gradually opened to maintain a constant manifold pressure until the high ratio full throttle, or critical, altitude has been reached. Above this, operation will be a t full throttle, with gradually decreasing manifold pressure. If in continuous level flight the pilot has occasion to alter his type of operation, this may sometimes require a shift in impeller ratios even though the altitude remains unchanged. For example : assume the pilot is flying a t 12,000 ft., a t Normal Rated power and 2600 rpm. As shown in the power plant chart, this calls for operation in the high impeller ratio ("HIGH") a t part throttle with a manifold pressure of 40.0 in. Presently the pilot finds a reduction in power is indicated, and he decides to operate a t Maximum Cruise power and 2250 rpm. Reference to the chart will show that these new conditions call for operation in the low impeller ratio ("LOW"), a t full throttle and 2250 rpm. Accordingly, the pilot must shift from "HIGH" to "LOW". A wider range of operating conditions is covered in more detailed charts. Where the particu-

lar type of, operation desired is not specifically dealt with on the check charts available, shift altitudes may be estimated by interpolation o r more accurately calculated from the appropriate operating curves as described in the previous section. The penalties for failure to shift at the proper altitudes have already been discussed. Determination by Rule of Thumb.-A rough rule for determining shift altitude is : 1. After full throttle operation in the low gear ratio at the selected rpm and manifold pressure has been reached, and has been continued until the manifold pressure has dropped 3 to 8 in. hg. (depending on the type of engine and the performance range), shift to the high gear ratio. 2. Readjust manifold pressure to the desired value. This is a t best a rough rule, and its correct and effective application depends to a large extent on the pilot's familiarity with his engine and its installation, as well as with the limiting rpm's and manifold pressures under which the engine may be operated.

1

I

f

PRATT & WHITNEY AIRCRAFT D I V I S I O N

O F

U N I T E D

A I R C R A F T

C O R P O R A T I O N

E A S T HARTFORD 8. CONNECTICUT, U. S . A .

BWA. PART NO. 116936

PWA. 01.60

r THE USE OF OPERATING CURVES

DECEMBER, 1946

lNSTALLATlON ENGINEERING -

f

W H I T N E Y

P R A T T D i V I S I O N E A S T

OF

UNITED

H A R T F O R D

A I R C R A F T

AIRCRAFT

CORPORATION

C O N N E C T I C U T

NOTE: Except for the sample power curves opposite page 32 the curves which follow in the text are for illustrative purposes only. The results obtained from them may be considered as typical, but are not applicable to any particular engine.

PRATT

&

P W A . 01. G O

WHITNEY AIRCRAFT

INTRODUCTION Purpose

0

Operating curves, often referred to as power curves, are popularly considered one of the dullest and least glamorous aspects of aircraft engine operation, and not infrequently they are dismissed as being too academic, if not hopelessly obscure.. Nevertheless, an understanding of their construction and a knowledge of their application provide a useful and convenient means of analyzing and predicting engine performance. For example: in planning a flight, or while actually in the air, a pilot or flight engineer may have occasion to estimate the brakehorsepower (bhp) delivered by his engine, and he must be able to do so rapidly and with an accuracy sufficient for practical purposes. Some engines are equipped with torquemeters, and in such cases the operator can calculate his bhp directly by means of a simple formula: BHP = K X RPM Xtorquemeter pressure reading,

0

without reference t o manifold pressure, carburetor air temperature '(c.a.t.), altitude, or other factors save rpm and torquemeter pressure reading. For engines not so equipped bhp is usually computed by means of performance charts or graphs known as operating or power curves, developed from calibrations of the actual engines to which they apply. Operating curves are also essential to other types of calculation. For instance: a pilot may wish to compute the manifold pressure necessary to produce a certain bhp a t a given altitude and rpm; he may want to know how his manifold pressure and bhp will be affected by a change in rpm or altitude; or he may have occasion to find a critical altitude, an efficient rpm-manifold pressure combinatioq or the limits of safe engine operation. power curves are used in 'connection with 'the data obtained from the following four instruments:

1. Tachometer (engine crankshaft rpm), 2. Manifold Pressure Gauge (intake mani'fold pressure, or main stage supercharger rim pressure, absolute, in inches of mercury - in. hg), 3. Altimeter -(altitude in feet as calculated from atmospheric pressure - pressure altitude), 4. Carburetor Intake Air Thermometer jc.a.t., in deg. C. or deg. F.). In addition, certain assumptions are implied, such as: carburetor deck and exhaust back pressures equal to those called for at standard pressure altitudes; absolutely dry air; correct cylinder and oil temperatures; correct fuel and oil specifications and pressures; and proper engine condition. Selection

Three typical operating curves are shown opposite page 32. I t should be note4 at the very start that power curves are not interchangeable, and the operator must make certain that the one selected is the correct curve for his particular: 1. Engine (model number; propeller gear, impeller gear, and compression ratios; impeller diameter; fuel metering; etc.), 2. Grade of Fuel (91,100, 100/130 etc.), 3. Mixture Setting (auto rich, auto lean, etc.), 4. Impeller Ratio (neutral, high, low'- if operating a single-stage, two-speed, or a two-stage, two-speed engine). These data may be found in the upper righthand corner and elsewhere on the chart. Failure to use the correct curve, as well as to use it properly, will yield results that are misleading - if not disconcerting.

PWA. 0 1 . 6 0


SEA LEVEL CALIBRATION CURVE ALTITUDE CALIBRATION

SEA LEVEL CALI B R A T I O N

22 24 26 28 30 32 34 36 38 40 M A N l F O L D PRESSURE

Fig. 7

- IN. HG

- Construction

0

of Sea Level Calibration Curve

Construction

An operating curve is divided into two parts: a sea level calibration curve on the left, and an altitude calibration curve on the right. The sea level calibration of an engine requires the use of a dynamometer, or power rneasy-ing device; and the data obtained are recorded in the form of a graph whose coordinates are manifold pressure in in. hg (horizontal) and bhp (vertical). With the engine on the test stand, the operator sets the throttle and adjusts the load imposed by the dynamometer to permit the engine to turn a t some desired rpm, say 1800. The bhp is then calculated from the dynamometer reading, and the manifold pressure taken from the manifold pressure gauge. The values thus obtained, say 350 bhp and 23 in. hg, determine the location of point A in Fig. 1. The operator then opens the

throttle a certain amount, a t the same time keeping the rprn constant at 1800 by increasing the dynamometer load. Bhp and manifold pressure are again read as before, and the new set of values, say, 480 bhp and 30 in. hg, fix the location of a second point, B. This procedure is repeated until the throttle is fully open and no further power can be delivered a t 1800 rpm. The bhp and manifold pressure readings a t full throttle determine point C. The points are then connected, and the resultant curve, ABC, is a part throttle, constant rprn line (ending at the full throttle point, C) which shows the relation between manifold pressure and bhp at the given rpm. Once the shape and position of this line have been established, the operator can by means of it compute the engine's bhp in terms of manifold pressure, provided the rprn remains constant a t 1800, and other conditions are unchanged.

1

C

MAN IFOLD PRESSURE

- IN. HG

Fig. 2 - Completed Sea Level Calibration Curve (for "Sea Level" Engine)

Since variable pitch propellers permit operation over a wide range of powers a t any particular rpm, i t is necessary t o plot a series of different constant rpm lines t o cover all reasonable combinations of rpm, bhp, and manifold pressure. Each of these lines is constructed in the same manner as the one just described. When the calibration has been completed the result will be a sea level calibration curve not unlike the one shown above in Fig. 2. Discussion

If the calibration is made throughout at the best power mixture setting (the fuel-air ratio that gives the highest power obtainable a t a (.'.particular manifold pressure, airflow, or throttle the part throttle, constant rpm lines are

straight lines, originating a t 0 in. hg manifold pressure and negative bhp values determined by the engine's characteristics. In practice the calibrations are also made with an actual carburetor set in the automatic rich or automatic lean position. Since the fuel-air ratio is varied with power t o satisfy the engine's requirements, the constant rpm lines will often be slightly curved, the curvature depending on the metering characteristics of the carburetor and the effect of fuel-air ratio on power obtainable a t various manifold pressures and rpm's. The line ABCDE in Fig. 2 is drawn through the full throttle points that terminate each of the part throttle, constant rpm lines. It is called a full throttle line, and indicates the maximum manifold pressure and bhp that can be obtained in sea level operation a t any rpm covered by the curve.

P W A . OX. 6 0

PRATT & W H I T N E Y AIRCRAFT SEA LEVEL CALI BRATI ON

ALTITUDE CALI BRATION

0

Fig. 3 - Restricted Ranges; Computation of BHP from Sea Level Calibration Curve

Since full throttle operation a t sea level is normally restricted or prohibited for all but socalled "sea level" engines, because of limitations imposed by the engines' construction and selection of fuel, the upper parts of the sea level calibration curves are frequently omitted, or shown merely as a series of dashes, as illustrated in Fig. 3. The location of the full throttle points can then be determined only by extrapolation of the curves or from theoretical considerations. The lower .portions of the curves are also omitted when they fall outside the range of normal operation. Some sea level calibration curves include a short line in the upper right-hand corner indicating the maximum rprn and manifold pressure permissible for restricted operation (such as Take-Off or Military Power) a t sea level. Since the part throttle, constant rprn lines are straight, or nearly so, the sea level curve may be

used to illustrate an important relationship, which holds approximately true for engine operation a t any given pressure altitude, namely: bhp (plus a constant) varies directly as manifold pressure, provided rprn and other factors are maintained unchanged. Bhp also varies as rpm, if manifold pressure and other factors remain constant; but the variation is not linear. If rprn is increased in increments of 200, for example, the corresponding increases in bhp a t any fixed manifold pressure become successively smaller, as may be seen by. the relative crowding of the rprn lines in Fig. 3 a t the higher speeds. (In extreme cases, an increase in rprn may actually result in a loss of power.) The point to bear in mind, however, is that bhp is a function of both rprn and manifold pressure. Any given bhp can be obtained, within operating limits, by a virtually infinite number of rpm-manifold pressure combinations, as may be seen from a sea level calibra-

T H E U S E OF O P E R A T I N G C U R V E S

tion curve by extending any bhp line across the part throttle, constant rpm lines, and noting the manifold pressure corresponding to each intersection. Use of Sea Level Calibration Curve

@

Once we have completed the sea level calibration for a particular engine, we can dispense with the dynamometer, and calculate bhp a t sea level from the curve whenever rpm and manifold pressure are known. Thus: if the tachometer reads 2000 rpm, and the manifold pressure gauge 28 in. hg, the sea level bhp of the engine, for which the curve in Fig. 3 was made, may be found as follows: 1. Locate the intersection of the 2000 rpm and 28 in. hg manifold pressure lines a t A. 2. Project this intersection horizontally to the bhp axis at B. 3. The bhp will there be found to be 500. (Since point A is below and to the left of the fuIl throttle point, C, 500 bhp represents part throttle power.) Values for readings that do not fall exactly on the lines of the graph may be found by interpolation. Carburetor Air ~ e r n ~ e r a t u r(c.a.t.) e Correction

The calibration just outlined was presumably made a t 0 ft. altitude and under standard sea level

conditions of atmospheric pressure and temperature: 29.92 in. hg barometer and 15" C. (59" F.) c.a.t.. If the sea level calibration curve is used a t altitude, or under conditions other than standard sea level conditions, errors will be introduced in the calculations and the results will be incorrect. The maintenance of a constant sea level altitude presents no problems, and variations from the standard barometric pressure may, for practical purposes, be considered as accounted for in the manifold pressure gauge readings, while other factors, such as the effect of water vapor present in the air, may be neglected. Carburetor air temperatures, however, are not readily controlled and may depart considerably from the standard 15" C. (59" F.). Corrections for these variations can be made with sufficient accuracy as follows:

1. To the bhp obtained from the curve add 1%for each 6" C. (10" F.) of c.a.t. below standard temperature; or 2. Subtract 1% for each 6" C. (10" F.) above standard.

b

Thus: if the c.a.t. gauge in the foregoing example had read 33" C. (91" F.) - i.e. 18"C. (32" F.) above standard conditions - the bhp of 500 obtained from the curve should have beer1 reduced by 3% to give the actual bhp of 485. Had it read 3" C. (37" F.) - i.e. the bhp 12" C. (22" F.) below standard should have been increased by 2% to 510.

-

ALTITUDE CALIBRATION CURVE The operating curve is completed by the altitude caIibration curve which lies directly to the right of the sea level calibration curve. (See typical power curves opposite page 32.) Sea level and altitude caIibration curves differ basically, the former being essentially a part throttle, the latter a full throttle operation curve. Except in a limited way the two cannot be used independently of one another. Construction

In constructing the altitude calibration curve

we must, of course, make use of the same engine that was calibrated a t sea level. Let us assume that it is a "sea level" engine, unsupercharged or incorporating simply a single-stage, single-speed supercharger. Atmospheric conditions will be considered as standard throughout the calibration. (Corrections for variations in c.a.t. will be discussed in connection with the use of the completed curve.) Operation in best power mixture setting will also be taken for granted. The engine is first run a t sea level with the throttle wide open and the propeller governor

\

PRATT

&

SEA LEVEL CALIBRATION

'

PWA. 01. 6 0

WHITNEY AIRCRAFT ALTITUDE CALIBRATION

M P - I N . HG

-

PRELIM. FORM

-

ALTITUDE I000 f T LINEAR SCALE

Fig. 4 - Construction of Altitude Calibration Curve - Preliminary

set to maintain an rpm of, say, 1800. Referring to the sea level calibration curve in Fig. 4, we locate the full throttle point, A, on the 1800 rpm line, and from its position find the manifold pressure and bhp to be 36 in. hg and 600 bhp respectively. Then, projecting the second of these values (600 bhp) horizontally, we establish the initial point, B, of the altitude calibration curve on the vertical axis, which is identical with that of the bhp axis of the sea level calibration curve. The manifold pressure a t B is shown by a short diagonal line. Since it lies on the 0 ft. altitude line, it has the same value (36 is. hg) as that shown on the sea level calibration curve. The engine is next taken to altitude, still a t f ~ h throttle l and with the speed kept constant a t the original setting of 1800 rpm. As we climb, the pressure and density of the atmosphere become progressively less, and, as the mass of air available to the carburetor decreases, both manifold pressure and bhp fall off. The relations between bhp, altitude, and manifold pressure are shown

by the graph in Fig. 4, where bhp is plotted against an altitude scale in feet, and manifold pressures are marked off along the curve. This 'curve is based on flight calibrations of many different types of engines, together with a number of experiments made on test stands where altitude conditions can be simulated. These tests show that, if full throttle, constant rpm operation is maintained, all engines will develop approximately the same percentage of their full throttle, sea level bhp a t any given altitude (with the exception of engines which make use of variable speed impellers or turbosuperchargers). For example: all engines (with the exceptions noted) develop approxilnately 75% of their sea level power a t an altitude of 8500 ft., and 50% a t 19,000 ft. Referring again to Fig. 4, we find that our engine delivers 440 bhp a t 9000 ft., a t full throttle and 1800 rpm, and that the manifold pressure is close to 25 in. hg. We also note that the bhp has decreased to 50% of its sea level value a t a n altitude of 19,000 ft.

THE USE OF OPERATING CURVES

n

SEA L E V E L CALIBRATION

M P - I N . HG

ALTITUDE CALI B R A T l O N

1.0

0:9

o;?

0:s

0.6

0.5

ALTITUDE-DENSITY RATIO LINEAR SCALE

Fig. 5

- Construction of

-

Altitude Calibration Curve Conversion of Linear Altitude-Density Scale to Non-Linear Altitude Scale in Feet

Now, engine power depends on the density of the air at the carburetor inlet; and extensive tests have shown that, at any fixed throttle setting (such as full throttle) and constant rpm, bhp may be considered for all practical purposes as varying linearly with the altitude-density ratio. (Altitude-density ratio, c,is defined by the equation: a=P /Po, where P is the density of the air a t a given altitude, A, and Po the density a t sea level.) Accordingly, if bhp is plotted against an altitude-density ratio scale, instead of an altitude scale in feet, the full throttle, constant rpm curve becomes a straight line, thus simpli-

fying its construction. For convenience altitude is still expressed in feet, whieh may be readily done by locating each altitude line opposite its corresponding density ratio, as illustrated in Fig. 5 above, where the relation between the two scales is shown. Since this relation is not linear, the altitude lines are unevenly spaced, becoming progressively closer together toward the righthand side of the curve. An altitude scale, based on the density ratio but expressed in terms of feet, is thus established for the second (horizontal) coordinate of our altitude curve.

PWA. 0 1 . 60

PRATT & WHITNEY AIRCRAFT SEA LEVEL C A L I B R A T I O N

MANIFOLD PRESSURE

ALTITUDE CALIBRATION

- IN.HG

Fig. 6

- Completed

ALTITUDE

- 1000 FT

Altitude Calibration Curve (for "Sea Level" Engine)

The tests previously referred to establish the fact that an aircraft engine delivers 50% of its full throttle, sea level bhp a t an altitude of approximately 19,000 ft., so long as the full throttle setting and rpm remain unchanged. Making use of this relationship we can proceed to draw all the necessary full throttle constant rpm lines with virtually no further calibrations. From the full throttle line ABCDE, of the sea level calibration curve we can determine the engine's full throttle, sea level bhp for as many different rpm's as we choose. Projecting these bhp values horizontally on to the sea level line of the altitude calibration curve, as before, we there locate the initial points, F,G,H,I,J, of the corresponding full throttle, constant rpm lines. Since the bhp a t 19,000 ft. is one-half of its sea level value (provided full throttle setting and rpm remain unchanged), if we take a series of bhp values a t 19,000 f t . equal, respectively, to 50% of the corresponding sea level values - in other words,

make PK- 54 OF; PL = OG; etc. - we have established a second point for each of the full throttle, constant rprn lines. Since the latter are straight lines (when plotted on the altitude-density ratio scale), we now have all the data necessary to determine their positions, and may forthwith draw the full throttle lines as shown in Fig. 6 above. The altitude calibration curve is then completed by adding a number of full throttle, constant manifold pressure lines, generally shown as a series of dashes drawn diagonally across the full throttle, constant rpm lines. Their positions can be established in several ways, including simple flight calibrations, since the readings of the manifold pressure gauge, the tachometer, carburetor deck pressure, and altimeter (with the -throttle in the full open position) alone are involved. The manifold pressures will be found to follow roughly the same percentages of their sea level values as do the bhp's.

I

THE USE O F OPERATING CURVES SEA LEVEL CALI ERATION,

ALTITUDE CALIERATION

2 MP- IN. HG

Fig. 7 - Restricted

0

I

Discussion

-

I

I

0

- 1000 FT - .. Ranges, .

ALTITUDE

From the geometrical construction of the full throttle, constant rpm lines we infer that they converge a t some point on or near the zero bhp, horizontal axis of the altitude calibration curve. This point appears to correspond to an altitude of about 55,400 'ft., where presumably all the power developed in the cylinders is'used to overcome .engine friction and drive the geared supercharger, thus reducing the bhp to nothing. This would seem to place a limit on the altitude which can be reached with a single or multiple stage, gear-driven supercharger. On the other hand, data taken from flights made at extreme altitudes suggest that the lower ends of the lines curve to meet the horizontal axis a t points above 55,400 ft., and it. is interesting to note that the world's altitude record is actually above this figure. One of the assumptions underlying the straight line construction of the altitude calibration curve is continuous operation in best power mixture setting. In practice this condition is seldom realized, and most altitude calibration curves are constructed for some actual carburetor setting, such as automatic rich or automatic lean. Where

'

actual Right caJ.i-bratiohsare available for such settings, the results indicate slight, but appreciable departbes frdmbthe'straight line geometrical constructions described. . As will be diicu&ed later, eiigine operation is normally lim'ited by one or more of the following factors: bhp, rpm, manifold pressure, or cylinder pressure, the last expressed indirectly in terms of the brake mean effective pressure (bmep). Again, just as in the case of 'the sea level curve, those parts of the alt?tude curire which lie beyond the range of normal operation are generally omitted; and the operatot's attention is called to restricted or prohibited ranges by dash lines or printed instructions. For example: on the power curve in Fig. 7. maximiim power available for 5 minutes only is limited'to 750 bhp and 2400 rpm, with manifold pi-essure's restricted accordingly to 34 in. hg at sea Ievel and 32.75 in. hg a t 5400 ft. Maximum continuous rating is similarly limited to 650 bhp and 2200 rpm, with manifold pressures restricted to 32 in. hg a t sea level and 30.75 in. hg a t 6400 ft.; and continuous full throttle operation is prohibited in the range above and to the left of the heavy line AB, which line represents operation a t maximum permissible cylinder pressures. 1

?.,


PWA. 0 1 . 60

Fig. 8 - Composite Calibrations of Single-Stage Single-Speed, Single-Stage Two-Speed, and Two-Stage Two-Speed Engines

In the case of "altitude engines", which cannot be operated a t sea level a t full throttle, the construction of the altitude calibration curve follows substantially the same procedure as that just outlined, except that the initial full throttle, sea level points must be determined by extrapolation or from theoretical considerations. (The typi'cal power curves shown opposite page 32 are for an altitude engine.) Had we been calibrating either a single-stage, two-speed engine, or a two-stage, two-speed engine, a separate set of operating or power curves (both sea level and alti'tude) would have been . necessary for each supercharging combination two sets of curves for the first engine, three for the second. Since the pumping capacity of an impeller increases with its speed, it follows that operation in a high impeller gear ratio results in a higher manifoId pressure and bhp than operation a t the same throttle setting and ,rpm in a low impeller gear ratio. The same result is obtained by running a two-stage engine with the auxiliary stage aiding the main stage at constant rpm's and throttle settings. This effect is illustrated on -the power curve by a shift of the rpm and manifold pressure lines to the right, as may be seen in the composite altitude calibration curve in Fig. 8.

I t should be noted that the gain in bhp with the added supercharging is not proportional to the increase in manifold pressure. This is because a greater fraction of the engine's power is absorbed in driving the impeller in the higher gear ratios, or in driving the two impellers of the two-stage system, and therefore a correspondingly smaller fraction of the total power develope'd is 'available as bhp a t the propeller shaft. I t should also be noted that all full throttle, constant rpm lines, regardless of impeller gear ratio (or stage) appear to converge at approximately the same 0 bhp, 55,400 ft. altitude point. This is true only of geared superchargers. Operating curves of the conventional type cannot be drawn satisfactorily to cover variable speed impeller or turbosupercharger operation, since the ratio of the speed of the, impeller to that of the crankshaft is not fixed.

.

Part Throttle Operation Before proceeding to an explanation of the use of the operating curve, we must first investigate the method of plotting part throttle operation on the altitude calibration curve. Returning to our sea level, unsupercharged (or single-stage, single-speed) engine, let us run it a t sea level a t an rpm of 2200 and a manifold pres-

I

J

THE USE OF OPERATING CURVES SEA LEVEL C A L I B R A T I O N

MP-

A L T I T U D E CALIBRATION

IN.HG

ALTITUDE

- 1000

FT

Fig. 9 - Part Throttle Operation of Altitude

0

0

sure of 24 in. hg. On the sea level calibration curve in Fig. 9 we locate the intersection, A, of the given rprn (2200) and manifold pressure (24 in. hg) lines. Projecting this intersection, A, horizontally to B and C, we note that this particular combination of rprn and manifold pressure produces 460 bhp a t sea level, and represents a part throttle operation. After setting the propeller governor to keep the rprn unchanged a t 2200, we then take the engine aloft, opening the throttle progressively as we climb t o maintain the manifold pressure constant a t 24 in. hg. As we ascend, without varying either rprn or manifold pressure, there will be a gradual increase in bhp, due chiefly to two factors: a decrease in the atmosph'eric back pressure on the. exhaust and a lowering of 'the c.a.t.. The first has the effect of emptying the cylinders more com pletely of exhaust gases and of reducing friction the second increases the weight of a given volume of the combustion charge a t the manifold pressure of 24 in. hg. Bhp will therefore continue t o increase until an altitude is attained a t which the throttle can be opened no further - in other words, when full throttle operation a t 2200 rpm and 24 in. hg has been reached. This point is readily determined from the altitude calibration curve by locating the intersection, D, of the full

throttle, constant 2200 rprn line and the 24 in. hg manifold pressure line. Projecting this intersection, D, horizontally to E, we find the hhp to be 530 (an increase of 70 bhp over the sea level value of 460); projecting it vertically downward to F, we find we have reached an altitude of 12,000 ft. This altitude, a t which the throttle is wide open and the available bhp is a maximum, is known as the critical altitude for the particular combination of rprn and manifold pressure. We have thus determined two points on the part throttle, constant rpm, constant manifold pressure line: one a t C from the sea level curve, the second a t D from the altitude calibration curve. For all practical purposes it can be assumed that the bhp, a t constant rprn and manifold pressure, increases uniformly on the altitudedensity ratio scale used. Accordingly we can drav the part throttle; constant rpm, constant mani fold pressure line, by simply connecting C and D with a straight line. Any point on the line, CD, will show the part throttle bhp available for the given rprn and manifold pressure combination a t the altitude of the point. For example: t o find the bhp available a t 5000 ft. for the given combination of 2200 rprn and 24 in. hg manifold pressure, locate the intersection, G, of the 5000 ft. altitude line and the constant rpm, constant

PWA. 0 1 . 60

PRATT & WHITNEY AIRCRAFT SEA LEVEL CALIBRATION

ALTITUDE CALIBRATION mnn

MP

- IN. H G

ALTITUDE

-

1000 FT

Fig. I 0 - Part Throttle Operation - RPM-MP-BHP Combinations

manifold pressure line, CD, and project this intersection horizontalIy to the vertical axis a t H, where the bhp is read as 490. Since G lies below and to the left of the full throttle point, D, part throttle operation is indicated. We shall do well at this juncture to establish clearly in mind the distinction between the following lines by referring to Figs. 9 or 10 above: (1) Part Throttle, Constant Rpm (Sea Level Calibration Curve), (2) Full Throttle, Constant Rpm (Altitude Calibration Curve), . (3) Part Throttle, Constant Rpm, Constant Manifold Pr-wure (Construction on Altitude Calibration Curve), (4) Full Throttle, Constant Manifold Pressure (Altitude Calibration Curve). A study of Fig. 10 above will show that point G actually lies on an infinite number of part throttle, constant rpm, constant manifold pressure lines. For example: it 'may be seen from the sea level calibration curve that 460 bhp is produced by a combination of 2000 rprn and 26 in. hg manifold pressure, by another of 1600 rprn and 32.5 in. hg, etc. as well as the one originally selected of 2200 rprn and 24 in. hg.

.

Full throttle operation a t the comblnatlon ot 2000 rpm and 26 in. hg manifold pressure determines the location of point I; and, if the line CI is drawn, it will be found to coincide with the line CD previously drawn. In the same way G may be found to lie on all the part throttle, constant rpm, constant manifold pressure lines that have their origin .at 460 bhp at sea level. (In practice, using actual power curves, the coincidence of these part throttle, constant rpm, constant manifold pressure lines may not be exact, but it is sufficiently close for all purposes.) This is simply another way of saying that all combinations of rprn and manifold pressure which deliver equal bhp's at dne altitude (such as sea level) will also deliver: equal (but not the same) bhp's at any other, provided operation is in the part throttle range. From the geometrical construction of the operating curve it can be shown that all part throttle, constant rpm, constant manifold pressure lines have practically the same dopes. (See lines JK, CI, and LM, in the Fig. 10 above.) The fact that these lines are all approximately parallel is a useful relationship in checking work done an the altitude calibration curves, and is basic to certain of the computations. (FulI throttle, constant rprn

THE USE O F O P E R A T I N G C U R V E S SEA LEVEL CALIBRATION


ALTITUDE

M P - IN. HG

Fig. I I

O

10

- Part

Throttle Operation

lines are generally not straight where altitude calibration curves have been plotted from data obtained with mixtures other than best yower, and full throttle manifold pressure lines may be decidedly curved because of variations in supercharger efficiency. Under any of these conditions the part throttle lines may not be closely parallel to one another throughout the entire range of the operating curve.) Another type of part throttle operation is represented by the constant bhp, constant rprn lines, such as the line AB in Fig. 11, where the bhp is 550 and the rprn 2400. Since engine performance is often limited in terms of these two quantities, these lines are of common occurrence on altitude . calibration curves. The critical altitude for the given bhp-rpm combination is 14000 ft., determined by the intersection, B, of the constant 550 bhp, constant 2400 rprn line with the full throttle, constant 2400 rprn line. Because of greater exhaust back pressure and higher c.a.t., the manifold pressure necessary to develop a given bhp a t a particuIar rpni will be greatest a t sea level, and will decrease uniformly with the altitude-density ratio until the critical altitude is reached. The manifold pressure a t sea level is found from the sea level calibration curve,

.

- 1000 FT

- Constant BHP-RPM

Lines

and proves in our example to be 26 in. hg; a t the critical altitude it is found from the altitude caIibration curve, and proves to be 23 in. hg - a decrease of 3 in. hg. If the constant bhp, constant rprn line, AB, is divided into three equal parts, each division will represent a 1 -in. hg difference in manifold pressure, and the altitudes corresponding to 25 in. hg and 24 in. hg manifold pressure turn out to be approximately 4200 and 8800 ft., respectively. (The fact that the manifold' pressure decreases uniformly from sea level to the critical altitude will also be evident from the geometrical construction of the curve, since the part throttle, constant rpm, constant manifold pressure lines, AD, CB, etc., are parallel, and the manifold pressure divisions along the full throttle, constant 2400 rprn line are equal. Where the lines of the sea level or altitude calibration are curved, these reIationships are only approximate.) If we wish to find the manifold pressure necessary to deliver 550 bhp a t 2400 rprn at an altitude of 2000 ft., we locate the intersection, E, of the given' altitude (2000 ft.) and bhp (550) lines, and note that it .lies approximateIy midway between the 26 in. and 25 in. hg manifold pressure points, thus indicating that a manifold pressure of about 25.5 in; hg is called for.

PRATT

&

WHITNEY AIRCRAFT

SEA L E V E L C A L I B R A T I O N

PWA. 0 1 . 6 0 ALTITUDE CALI 6 RATION

'

.

24 26 28 30 32 34 36 38 40 MP- IN. HG

Fig. 12 - Given: BHP

1

The Use of the Complete Operating or Power Curve I

The problems which involve the use of operating curves fall generally into one of.two types: A. The calculation of manifold pressure, when bhp, rpm, and altitude are known; B. The calculation of bhp, when manifold pressure, rpm, and altitude are known. Examples of these problems are given below: T g p e A -First Solution : (See Fig. 12.) A. Given: Bhp = 480 To Find : ManiRpm =2200 fold Pressure Alt. =8000ft. 1. Locate the intersection, A, of the given altitude (8000 ft.) and bhp (480) lines. 2. Select anyconstant manifold pressure line (say, 26 in. hg) estimated to be close to the required answer, and locate its intersection, B, with the full throttle, 'constant rpm line corresponding to the given (2200) rprn line. 3. Transfer these vaiues (26 in. hg manifold pressure a.nd 2200 rpm) to the sea level calibration curve, and locate the intersection, C, of the 26 in. hg manifold pressure and 2200 rpm lines.

4. Project this intersection, C, horizontally to the bhp axis of the altitude calibration curve, and locate the point D. 5. Connect D and B by a straight line, DB. 6. Through point A draw a line, EF, parallel to line DB. 7. Locate the intersection, G, of the line E F and the full throttle constant 2200 rpm line. The required manifold pressure (23 in. hg) will be found a t point G.

IVOTE: Since point A lies below and to the left of the full throttle, constant 2200 rprn line, part throttle operation is indicated. Explanation : Since all part throttle, constant rpm, constant manifold pressure lines are approximately parallel, it follows that, if we determine the slope of one such line, DB, we can draw the corresponding line, EF, through the given bhp-altitude point, A. Inasmuch as manifold pressure is constant a t all points on the line, EF, its value may be found a t that point, G, where the part throttle, constant rpm, constant manifold pressure line, intersects (or, better, terminates in) the full throt-

THE USE O F OPERATING CURVES SEA LEVEL CALIBRATION

M P - IN. H G Fig. 13 - Given: BHP

ALTITUDE CALI BRATI ON

=4

tle, constant given rprn line. The more closely we can estimate the desired manifold pressure, the more accurate wiIl be our construction; and in this case experience might have suggested that we use 24 in. hg for our preliminary estimate of manifold pressure instead of 26 in. Type A

- Second

Solution: (See Fig. 13.) Given: Bhp = 480 To Find: NIaniRpm =2200 fold Presswe Alt. =8000 ft.

1. Locate the intersection, A, of the given altitude (8000 ft.) and bhp (480) lines. 2. Determine as closely as possible the' full throttle, constant rprn line (1850 rpm) and the constant manifold pressure line (26.5 in. hg) on which this intersection, A, lies. 3. Transfer these values (1850 rprn and 26.5 in. hg manifold pressure) to the sea level curve, and locate the intersection, B, of the 1850 rprn and 26.5 in. hg manifold pressure lines. 4. Project this intersection, B, horizontally until it intersects the given (2200) rprn line a t C. 5. Project this last intersection, C, vertically

to the manifold pressure axis, and locate point D. The approximate required manifold pressure (23 in. hg) .will be found at

The intersection of t.he given bhp and altitude lines a t A shows that 480 bhp can be delivered a t 8000 ft. a t full throttle, 1850 rpm, and 26.5 in. hg manifold pressure. From the sea Ievel calibration curve we find that this combination of rprn and manifold produces a hhp a t sea level of 430 bhp. This same 430 bhp can be obtained a t sea level by all the rprn manifold pressure combinations that lie along the line BC extended, including the combination 2200 rprn and 23 in. hg manifold pressure (i.e. the given rprn and the manifold pressure necessary to produce 430 bhp at sea level). Since any two combinations of rpm and manifold pressure that give equal bhp's a t sea level will give approximately equal (though not 'the same) bhp's a t altitude, it follows that, if 1850 rprn and 26.5 in. hg manifold pressure, and'2200 rprn and 23 in. hg manifold pressure both give 430 bhp a t sea level, they will both give 480 bkp a t 8000 ft.

PRATT & W H I T N E Y AIRCRAFT

P W A . 01. 6 0


Fig. 14 - Given: MP = 26 in. Hg., RPM

Of the tw.o solutions, the first is probably the more accurate, if the full throttle, constant rpm and full throttle, constant manifold pressure lines are markedly curved, provided fairly accurate initial estimates can be made; the second solution possesses the ad. vantage of simplicity, and can be performed without any drawing instruments. The two methods offer convenient checks on one another, and the second may be used to advantage in making a preliminary estimate of required manifold pressure. Results obtained from power curves are generally con. siderkd to be accurate to f 2%%, if all work is carefully done. T y p e B - Solution: (See Fig. 14.)

Altitude = 4000 ft. l . * ~ o c a \ ethe intersection, A, of the given rpm (2000) and manifold pressure (26 in. hg) lines on the sea level calibration curve. 2. Project this intersection, A, horizontally to the bhp axis of the altitude calibration curve, and locate point B.

= 2000, AlT. = 4000

Ft.

Find: BHP

3. Locate 'the intersection, C, of the full throttle, constant rpm line and the full throttle constant manifold pressure line, corresponding respectively to the given rpm (2000) and manifold pressure (26 in.

4. Connect B and C with a straight line, BC. 5. Locate the intersection, D, of the line' BC with the given altitude (4000 ft.) line. 6. Project this intersection, D, horizontally to the bhp axis, and locate point E. The required bhp (485) will be found at point E.

NOTE: Since D lies below and to the left of the full throttle, constant rpm line for the given.rpm (2000), part throttle operation is indicated.

.

The engine's sea level bhp a t the given combination of rpm and manifold pressure is found from the sea Ievel calibration curve a t B. The bhp a t the full throttle critical altitude for the same combination is found from the altitude calibration curve at C. The line BC is therefore the part throttle,

r

THE USE OF OPERATING CURVES

constant rpm, constant manifold pressure line for the given combination of rpm and manifold pressure. The bhp for any altitude between sea level and critical altitude is then determined by the location of the intersection of the given altitude line with the part throttle line, BC. Carburetor Air Temperature Correction :

0

As the result of extensive investigations of atmospheric conditions a t sea level and a t altitude, a set of "standard conditions" of pressure and temperature have been established; and all operating curves are constructed on the assumption of the existence of these conditions. Departure from standard pressures, while they may result in false altitude readings in terms of number of feet above sea level, do not affect the use of the power curves. The airplane's altimeter when indexed to 29.92 in. hg (rather than to the altitude of the field) records altitude on the basis of pressure, and it makes no difference, as far as engine performance is concerned, whether a pressure of 19 in. hg is encountered a t 12,000 ft. (where it should be under standard conditions), or a t 11,000 ft. - the altimeter in either event will register 12,000 ft., and the curves will indicate performance as of that altitude. On the other hand, departures of c.a.t. from standard must be corrected for, since they affect the density and, hence, the weight of the combustion charge a t a given pressure altitude. Standard altitude temperatures may be found in atmosphere tables; they are also shown by means of the heavy Iine printed along the lower part of most altitude calibration curves. (Referring to any of the typical power curves opposite page 32, it will be seen that standard temperature a t 4000 ft. is 7" C. (45" F.); a t 8000 f t . is --loC. (30" F.); a t 40000 ft. is -55" C. (-67" F.); etc.) . Corrections for variations in c.a.t. as they affect bhp can be made with sufficient accuracy as before, namely : 1. To the bhp obtained from the curve add lyofor each 6" C. (10" F.) of c.a.t. below standard temperature; or

2. Subtract 1% for each 6" C. (10" F.) above standard. Thus: if the c.a.t. gauge in example B (Fig. 14) had read 19" C. (66" F.) - i.e. 12" C. (21" F.) above standard conditions the bhp of 485 obtained from the curve should have been reduced by 2% to give the actual bhp of 475. Had it read -11" C. (12" F.) - i.e. 18" C. (33" F.) below standard - the bhp should have been increased by 3y0 t o 500. -

Since bhp is a function of manifold pressure, the latter must be corrected for c.a.t. variations in problems which involve computing the manifold pressure necessary to produce a given bhp. The weight of the fuel-air charge delivered to the cylinders in a given time controls the engine's power output; and, if the temperature of the charge is raised, its weight per unit volume at a particular manifold pressure decreases and power output falls off correspondingly. Hence, to maintain the weight of the charge and the engine's bhp, manifold pressure must be increased as the c.a.t. is raised, and, conversely, decreased as the c.a.t. is lowered. Corrections for variations in c.a.t. as they affect manifold pressure can be made with sufficient accuracy as follows :

1. T o the manifold pressure obtained from the curve add in. hg for each 6" C. (10" F.) of c.a.t. above standard temperature; or 2. Subtract jd in. for each 6" C. (10" F.) below standard. Thus: if the c.a.t. gauge in example A (Fig. 12 or 13) had read 5" C. (41" F.) - i.e. 6" C. (11"F.) above standard conditions the manifold pressure of 26.5 in. hg obtained from the curve should have been increased by in. hg to give the actual manifold pressure required of 26.75 in. hg. Had i t read -24" C. (-11" F.) - i.e. 23" C. (41" F.) below standard - the manifold pressure should have been decreased by 1in. hg to 25.5 in. hg.

PRATT

&

WHITNEY AIRCRAFT

PWA. 01. 6 0



M P - IN.HG

Fig. 15 - Effect of Ram

Effect of Ram

The carburetor air intake scoop is usually faced into the slipstream in such a way as to gain a supercharging effect from the velocity pressure of the air stream. In a well designed installation this will provide -a considerable pressure rise, or ram, a t the carburetor deck. The result is t o increase the altitude a t which full throttle operation is reached for any given combination of rpm and manifold pressure - in other words t o increase the critical altitude. The lower exhaust back pressure and c.a.t. a t the new critical altitude result in additional bhp which can be computed from the power curve as shown in Fig. 15: Let us assume that flight data show sufficient ram pressure can be abtained to permit the engine t o operate a t full throttle, 2200 rpm, and 30 in. hg manifold pressure a t 8400 ft. instead of a t 6600 ft. (point A), the critical aICitude without ram. The horsepower with ram pressures may be found from the curve by constructing the part throttle, constant rpm, constant manifold pressure

line, BA, for 2200 rpm and 30 in. hg, and then extending the part throttle line, BA, until i t intersects, a t C, the line for the critical altitude with ram (8400 ft.). Projecting this intersection, C, horizontally t,o the bhp axis, we find the bhp to be 655 a t D - an increase of 15 bhp over that obtainable without ram for the same rpm and manifold pressure combination. Bhp for altitudes below the critical altitude with ram may, for practical purposes, be determined as before from .the intersection of the altitude lines with the part throttle, constant rpm, constant manifold pressure line, BC. Corrections for c.a.t. should be made a t the new critical altitude. 'The amount of ram depends on airspeed and on the design of the installation. It will therefore be different for each type of airplane and for every type of operation (climb, cruise, Military Power, etc.). Since it is impractical t o make charts t o meet all of these conditions, power curves are constructed for operation without ram, and the effect of the latter is computed as above after the relevant data have been obtained.

THE U S E OF OPERATING C U R V E S

SEA L E V E L CALIBRATION

0

24 26 28 30 32 34 36 38 40 M P IN. HG

-


0

2

4 6 PILTITUDE

-

8 10 1000 FT

12

14

I6

18

20

Fig. 16 - Critical Altitudes

Critical Altitudes and Engine Ratings

Y

I

0

Critical altitude has been defined above as the altitude a t which the maximum bhp is obtained with a given combination of manifold pressure and rpm; or,, stated another way, it is the maximum altitude attainable with a given manifold pressure and rprn combination. Full throttle operation is necessarily implied a t the critical altitude, since any possible further throttle opening would permit an increase in bhp and altitude. For example: if operated a t 32 in. hg manifold pressure and 2400 rpm, the critical altitude for our engine is 6000 ft., as illustrated by the solid red lines in Fig. 16 above. I t is also common practice to express critical altitude as the maximum altitude which can be attained with a given combination of bhp and

rpm. This can readily be found from the altitude calibration curve by locating the intersection of the given bhp line and the given full throttle, constant rprn line, since full throttle operation is likewise implied at this critical altitude. For example: if operated at 600 bhp and 2200 rpm, the critical altitude for our engine is 8600 ft., as shown by the red dash lines in Fig. 16 above. In this instance the manifold pressures are those required to give the desired bhp-rpm combination. They are naturally somewhat higher a t sea level (30 in. hg) than at. the critical altitude (27.75 in. hg) because of the greater exhaust back pressures and higher c.a.t's. encountered a t sea level. An engine's "critic81 altitude" is customarily understood to be the one attained when operating at the rated limits of rprn and bhp or manifold pressure.

PRATT

(I-.

&

WHITNEY AIRCRAFT

PWA. 0 1 . 6 0

.

ALTITUDE CALI BRATI ON


M P - IN. HG

ALTtTUDE

- 1000 F T

Fig. 17 - Ratings and Restrictions

Most supercharged engines may be operated a t or near sea level a t part throttle only, since the cylinder pressures and bhp's resulting from the use of the maximum manifold pressures available a t full thro!Ale are likely to prove destructive. Engine speeds must likewise be held below the point a t which dangerous strains are imposed by excessive rpm's. While it is sometimes conyenient to limit engine operation in terms of manifold pressure and rpm, it is customary to restrict it by holding cylinder pressures below a maximum value and specifying certain limiting values for bhp and rpm. Manifold pressures in themselves are not harmful to the engine, and we may let them come as they will, provided the bhp's, rpm's, and cylinder pressures to which they give rise are kept within the proper limits. Assume, for example, the limits for our engine are 650 bhp and 2400 rpm, with some maximum' permissible cylinder pressure understood. The restricted or prohibited ranges of operation are shown above in Fig. 17 by a series of dashes.

+

The shaded areas in the upper corners of the sea level and altitude calibration curves. represent bhp's theoretically available, but in excess of those which may be safely used for an extended period. Continuous full throttle operation is also prohibited in the range lying to the left ,of the limiting manifold pressure line AB, as'well as to the right of the corresponding sea level' line CD, since the manifold pressures called f6r will result in excessive cylinder pressures. The limiting bhp line, DBE, indicates part throttle operation from sea level to the critical altitude, with manifold pressures decreasing gradually from 30 in. hg to approximately 27.5 in. hg. Since the pilot is maintaining constant bhp, he will be opening his throttle less rapidly than. if.he were attempting to maintain constant' manifold pressure (as,for instance, along the line FE). It will also be noticed that certain combinations of bhp and rpm which are below the limiting, values of power and speed call for manifold pressures in excess of 30 in. hg (for example, a t points A and C.)

T-HE U S E O F O P E R A T I N G C U R V E S

- -- PART THROTTLE AT 2400 RPM 8 LlMlTlNG BHP FULL THROTTLE 2400 RPM

-- -

PART THROTTLE AT 2 4 0 0 RPM & L l M l T l N G B H P FULL THROTTLE AT 2400 R P M

-

0.

AT

a I m

I

m

SUPERCHARGl NG ALTITUDE TWO STAGE, TWO-SPEED

ALTITUDE SINGLE-STAGE, TWO-SPEED

Fig. 18 - Critical Altitudes: BHP-RPM

Had our engine been a single-stage, two-speed engine, or a two-stage, two-speed engine, we should have had one series of critical altitudes for each ,supercharger combination. If engine performance had been limited in terms of bhp and rpm (say, 600 bhp and 2400 rpm), with a maximum cylinder pressure implied or expressed, the composite altitude curves for the two types of engines just mentioned would resemble those in Fig. 18 above. The shaded areas represent the gain in power resulting from the additional supercharging. It will be noticed, however, that the maximum bhp available becomes less with each successive critical altitude. The total power output of an engine is designed to be approximately the same a t each altitude, but, since a greater prop6rtion of i t is absorbed in driving the impellers in the higher

gear ratios, or in driving the two impellers of the combined main and auxiliary stages, it follows that less bhp is available at the propeller shaft (A limiting mixture temperature may also call for a reduction in bhp.) If engine performance had been limited in terms of manifold pressure and rpm (say, 32 in. hg and 2400 rpm), the composite altitude curves for the two types of engines would have resembIed those in Pig. 19 below. These curves repeat the general characteristics of those in Fig: 18: a gain in power resulting from the additional supercharging, as shown by the shaded areas, with bhp's becoming less with each successive critical altitude because of power losses to the impeller. The location of points S indicate the proper altitudes a t which shifts should be made from one impeller ratio (or stage) to the next.

--- PART THROTTLE

-- - PART

--

THROTTLE AT 2 4 0 0 RPM, 32 I N . M P FULL TH-ROTTLE AT 2400 RPM

AT 2 4 0 0 RPM, 3 2 l N . M P FULL THROTTLE AT 2400 RPM

SUPE9CHARGI NG I

ALTITUDE SI NGLE-STAGE TWO-SPEED

,

.

I

ALTITUDE TWO-STAGE, TWO-SPEED

Fig. 19 - Critical Altitudes: MP-RPM

PRATT & W H I T N E Y AIRCRAFT

PWA. 0 1 . G O -


8000 F f CALIBRATION

t

1 i

16 I 8 20 22 24 26 28 30 32 34

0

2

4

6

Fig. 20

The complete operating curve consists basically of a part throttle and a full throttle curve.- The former is generally constructed from a calibration made at sea level, but, for various reasons, it may be desirable to make part throttle calibrations at altitudes other than this. Certain types of engine operation are of necestity limited to altitude, for example: operation in the high gear ratio of the auxiliary impeller -of a two-stage, two-speed engine. In such cases there is little to be gained by extending the altitude calibration below the operating range. At the same time the results obtained from the power curve will be more accurate if the part throttle calibration is made, not a t sea level, but a t some altitude corresponding more closely to the lowest part of the operating range, such as 8000, 10,000, or 12,000 feet. A typical 8000 ft. calibration curve is shown above a t the left-hand side of Fig. 20. I t is constructed in the same manner as the sea level calibration curve, except that the data have been obtained in flight at 8000 ft. The two curves are used in exactly the same way. The position of the 1600 rpm line of the sea level calibration curve is shown for purposes of comparison. I t will be noticed that the rpm lines have shifted upwards slightly, since the lower back pressures and c.a.tJs.

I

C

22

10 12 I000 F T

-

14

16

18

20

22

- 8bOO Ft. Calibration Operating Curve

Part Throttle Calibrations at Altitudes Other than Sea Level

I

8

ALTITUDE

M P - I N HG

-

at 8000 ft. result in higher bhp's for the same manifold pressures than a t sea level; the full throttle line has been displaced to the left, since the decreased atmospheric density a t 8000 ft. resuIts in lower full throttle bhp's and manifold pressures than a t sea level. The altitude calibration is not affected, except that it is cut off a t 8000 ft., the altitude at which the part throttle calibration was made, instead of being continued downward toward sea level. This omission is not significant, however, since virtually all of the curve below 8000 ft. would be i n t h e prohibited or restricted range. Within its range the altitude calibration curve is used in conjunction with the 8000 ft. calibration curve in the same manner it is used in conjunction with the sea level calibration curve, with the important exception that the bhp axis of the altitude calibration curve is now at 8000 ft., not at sea level. As a result the part throttle, constant rpm, constant manifold pressure lines slope upwards from the 8000 ft. line, rather than from the sea level line. (The problem of finding the bhp a t 12,000 ft., 1800 rpm, and 20 in. hg manifold pressure is solved in Fig. 20. It will be seen that the solution is of the same type as that in Fig. 14, except that the part throttle, constant rpm, constant manifoId pressure line originates at C, a t 8000 ft., and not a t B, a t sea level.)

1 i

4

THE USE O F O P E R A T I N G CURVES

BRAKE MEAN EFFECTIVE PRESSURE Engine operation is limited by a number of different factors, some of which have already been discussed. High rpnl's result in destructive centrifugal and reciprocating strains. Large bhp's impose dangerous loads on engine parts. High manifold pressures, while not in themselves harmful to the induction system, force into the cylinders heavy charges of fuel and air whose combustion may lead to excessive pressures and temperatures, if not to detonation. Closely reIated to these factors is another basic limitation: the maximum pressure which may be developed ,in the cylinders during the power stroke. This important limitation, which is dictated largely by engine construction, is normally understood in all engine ratings; when specifically stated it is usually expressed indirectly in terms of brake mean effective pressure (bmep) in psi. Cylinder .pressure is also a useful guide to the pilot in selecting the particular combination of rpm and manifold pressure best suited to a given type of operation. Any bhp may be obtained, as we have noted, by an infinite number of such

combinations, provided only it falls within the part throttle range. Since manifold pressure and oylinder pressure are closely related, this is equivalent to saying that a desired bhp may be obtained with an endless number of rpm-cylinder pressure combinations. On the one hand, if a low rpm is selected, the cylinder pressure must necessarily be high and may impose dangerous strains on the engine or bring about high cylinder head temperature that will result in detonation. On the other hand, if, to avoid this condition, a high rpm is selected, internal friction, wear, and,fuel consumption will all be increased, with consequent inefficient operation. Optimum performance will be found somewhere between these two extremes, at, or not far below, the maximum cylinder pressure which may be safely used for prolonged operation. Bmep proves to be a convenient,'if indirect, way of estimating these actual cylinder pressures and a nodding acquaintance with this often misunderstood and maligned term should assist the pilot in protecting his engine against abuse and in applying its power as efficiently as possible.

TOO LOW

TOO LOW

TOO HIGH

000 SPEED

PWA. 0 1 . 60


-COMPRESSION&

BOTTOM CENTER

POWER

EXHAUST

INTAKE

-

LLNTER

PI S T O N POSITION ( ~ Mt E OR CRANKSHAFT ROTAT ION)

Fig. 21

- Cylinder

Pressures during Power Cycle

Definition

Let us assume that we have a one cylinder, four stroke cycle engine turning at sonle given rpm. If the pressures present in the cylinder during a complete power cycle are plotted against piston position (expressed in terms of time or crankshaft rotation), the resulting graph will resemble the schematic pressure curve shown above in Fig. 21. The height of the curve above the horizontal axis represents the pressure in the cylinder a t any given instant. The shaded areas, P and I, below the curve represent power developed in the cylinder: P, by the combustion of the fuel-air charge; I, by the pressure of the incoming charge. The unshaded areas, C and E, represent the power expended in the cylinder: C, in compressing the charge; E, in expelling the exhaust gases against back pressure. The net useful power, or indicated horsepower (ihp), developed in the cylinder during one complete power cycle is therefore represented by the difference in area between the shaded and unshaded portions, or: IHP = (Area P +Area

I ) - (Area C +Area

E)

For many purposes it is convenient to conceive useful and of the various cylinder pressures adverse - as a single, uniform pressure acting during the power stroke alone, and capable of producing the given ihp. If we calculate the areas of the shaded and unshaded portZionsbelow the pressure curve, and determine the area which -

I

represents ihp, we can then draw a constant pressure line, AB, for the power stroke a t such distance above the horizontal axis that the area enclosed by the rectangle ABCD is equal to the area representing the ihp. In other words, draw AB to DC in such a way that:

a

.

Area ABCD = (Area P +Area I) - (Area C +Area E) = IHP

The mean, or average, power stroke pressure represented by the line AB is known as the indicated mean effective pressure (imep). Actually we are more concerned with the shape of the pressure curve than with the imep. However, experiments have established definite relationships between the two under a variety of different conditions; and, therefore, if we are able to measure the imep, we can estimate with sufficient accuracy the significant peak pressures and other combustion characteristics revealed by the configuration of the graph. Not all the power developed in the cylinders ii.e. the ihp) is available a t the propeller shaft. As shown in Fig. 22, some of it is lost in overcoming internal engine friction; some is diverted to driving the accessories; and some is absorbed in driving the supercharging mechanism. The power thus lost is called friction horsepower (fhp). What remains in the form of useful power delivered to the propeller shaft is known as brake horsepower (bhp). The mechanical efficiency of an engine is expressed as the ratio of power output to power input - i.e. ME = BHP/IHP.

I

I

0

T H E USE O F OPERATING CURVES

i i POWER STROKE

I H P = BHP

+F

IMEP= BMEP

+

-

-

. , .

a

-

'7

-. ,

,:#

a

,.-,-,F;A.h&,..'.

.

'

'

.

. I

%he dir&t compub,tion of ;an2:airh.#t*aigine'g * ':tb:e3hg:@9:bT&dily caloulated ffom the-qua.-. vti@$?+t.- ."w #(:-' , .. . . *, ihp-.gfid ;th+co.ngtr~ctionqi..ja,.~~esse;.,d>agra@~ ::-I- f y ~ - l $ ~ r ~ - q - u <- . i , , , ;.#m,~,&&+m.$. k&swSdnsia$bk ' d f i ~ ~ l t i d ir l . b ,%lie; gtk!r. - ;,j: hsnd ? ' &pi ca@,be.. *el,Q$&~j,:.eass~ spd &@&hp'i@ik 6&iehcy. cy.cy.f.~rn the.:forrnula. C.qei&;&ii.&l &,mai+cirij@tF,.;a@iLeue~+ei,. . . $; , - i i-IL~-;I~~t~2dw9iBrri;~;t;E:~jj.. . ;:. #'I ,#J. ,., - - .. " . ,"-, &~~+t+ h. e x..,~ $ i & i i,..p.$#&$2,ar;i;~Bi' ;~~ i3n$>fhppi,*hj$. -?*;. $ot. c*&*t7.' .:& are. em think bfii & e ~ ..be& . . .:. . - , . r;l.UUqtjt~, ,~$~:;.$&:fo@:~f~ A.5 , . :s;!, . . ... '.\ ::. - io .., p i & ~ ~ a l l ; . l r p ~ ; a ~ d ~. ....... ~ 9.,-..~ ..~. ., ~ ~., ..& g j ~ b &.&p-j, i f i.--3~ !. . .:.~-..b- ,$8,:y:.w&ea'd ,~...-,*.A. fie .. nieasurink, :od :a:rno$.$*≪lija~Om~ter, -,+,he. ?d%%ifi@r .to producing .bhp. -That,Ea+t. of the ppwer n&esWy 20 rnott:@th@e$&e rpithl-~l?,' '&
*..;:.

-

.

=-.mi

'-L:,l.,.

a

,,.I

.

,

.

.-.I,

+-

.

'

.

.

& *L

.

..

.

, ,

.

+;.

..,i-.

mp,

PRATT

&

WHITNEY AIRCRAFT

may be found by applying the following formula: 792000 BHP B M E P (psi) =-XEngine Displacement (cu. in.) R P M or :

BHP BMEP (psi) = K --RPM

P W A . 01. 60

6. The amount of work done per min. by an engine is equal to the work done per revolution times the rpm. Therefore: P XDispl. X R P M W/min. = 24

Where W/min. =no. of ft.-lbs. of work per min.

Bmep bears the same relation to imep that bhp does to ihp., namely:

7. Again by definition, power is the rate of doing work. (If you raise 50 lbs. 6 ft. in 2 sec., your power output will be four times as great as if you raise 50 lbs. 6 ft. in 8 sec.) Thus: Power = Work/Tirne

In other words, bmep may be defined as that portion of the imep which produces the given bhp. Derivation o f B M E P Formula

1. By definition, work is the product of a force by the distance through which it moves. (If you raise 50 lbs. 6 ft., you will do 300 ft.-lbs. of work.) Hence:

(It will be noted that the formula in (6) above is an expression of power.) 8. The conventional unit of power is the horsepower (hp). I t is defined as work done a t the rate of 33,000 ft.-lbs. per rnin. (If you weigh 165 lbs. and climb 200 ft. in 1 min., or 10 ft. in 3 sec., you will be putting out 1 hp.) Hence:

HP =

No. ft.-lbs. per min. 33,000

W=FXD Where: W = work in ft.-lbs. F =force in lbs. D =distance in feet.

9. I t follows, then, that the hp developed by our engine is given by the expression:

2. I n the case of en engine, the force exerted in each cylinder a t any instant is the product of the pressure a t that Ijarticular instant by the area of the piston head on which it acts. Thus:

HP=--

33,000

3. Assuming the cylinder pressure (and hence the force) remains constant through the power stroke, the distance moved by the force will be the length of the stroke. Accordingly, if P is the mean effective pressure: Where S =length of stroke in inches, divided by 12 to keep the distance in feet. 4. The work, in ft.-lbs., done per revolution by an engine of N cylinders is therefore:

Where N =no.'of cylinders, divided by 2 since the engine makes two complete revolutions during each power cycle (i.e. half the cylinders fire per revolution). 5. Since the total displacement of an engine is the product of the area of the piston head, A, by the stroke, S, by the number of cylinders, N, the formula in (4) above simplifies to: 12 x 2

24

Where Displ. =total engine displacement in cu. in.

-

P XDispl. XRPM 24 X 33,000

10. And, transposing : p=

Where: P =cylinder pressure in psi. A =area of piston in sq. in. (It will be noted that the sq. in's. cancel one another, leaving a force, F, in lbs.)

W/min.

792000XHP Displ. XRPM

=K -

HP RPM

Where K is a constant which depends on the displacement of the engine. For an R-2000 engine, for example: K =792000/2000 =396.

11. Since "P" has been assumed to be the imep,"HP." is by definition ihp. On the other hand, if "HP" 1s understood to be the bhp of the engine, "P" (the portion of the imep required to produce the bhp) becomes "BMEP" by definition. And so, finally: 792000 XBHP BHP = KBMEP (psi) = Displ. X R P M RPM (Note: Similar expressions may be used to compute frictlon mean effective pressure or other breakdowns of imep for analytical work.)

Discussion

Bmep, therefore, proves to be something of an abstraction - a fictitious pressure which has no real existence in the cylinders. Imep is clearly greater than bmep; and the actual peak pressures, as shown on pressure diagrams, are in turn greater than the imep. The abstraction, however, proves to be a useful one like "suction", an imaginary concept used to indicate a pressure differential. -

T H E U S E OF OPERATING CURVES

Since bmep may be readily calculated in terms of quantities that are easily found - bhp, rpm, and a constant - and inasmuch as it bears a definite relationship to imep and the latter, in turn, to peak pressures and other combustion characteristics, bmep serves as a convenient index - an indirect way of estimating actual cylinder pressures - and so provides a convenient scale by which to limit or rate engine performance. Since a rise in bmep will result in an increase in actual cylinder pressures, structural considerations will dictate operation below certain bmep values. Since high cylinder pressures will normally be accompanied by high cylinder head tefnperatures, it follows that bmep values set a series of limits for preventing detonation, for protecting the materials of the cylinders, and even for selecting a type of spark plug.

it can be seen that bmep varies directly as the bhp and inversely as the rpm. Accordingly, a pilot who drastically reduces his rpm without first reducing his bhp (by lowering his manifold pressure) runs the risk of increasing his bmep, and hence his actual cylinder pressures, beyond safe limits. Conversely, any substantial increase in bhp is dangerous unless preceded by a rise in rpm. In general it is desirable to run an engine a t the lowest possible rpm consistent with propeller efficiency, since fhp is thereby minimized, oil and fuel consumption diminished, and wear reduced. The most efficient results will therefore be obtained by operating at or near the maximum permissible bmep limits, since the higher the bmep, the greater the bhp for any given rpm - or, to put it another way, the maximum bhp will be secured with the minimum rpm. This type -of operation is particularly applicable to the cruising range, where no sudden increases in power which might result in exceeding safe bmep limits are anticipated.

From the relationship: BMEP =K-

BHP

RPM

NORMAL

BMEP

PROPERLY

AIR SPEED

RPM

BALANCED

PRATT & W H I T N E Y A I R C R A F T

PWA. 0 1 . 6 0 ALTITUDE CALIBRATION

SEA LEVEL CALI BRATI ON

.MP

- IN.HG

ALTITUDE:

- 1000 FT

Fig. 23 - Max. BMEP Cruising Curve

Bmep Curves

I

-v

d ''

N. :*:I

ikI ;

To aid the pilot in quickly selecting a combination of rprn and manifold pressure values which will result in operation' a t or just below the maximum permissible bmep, a series of curves known as Brnep Cruising Curves have been prepared. They are identical with the automatic lean curves, except that they show the part throttle manifold pressures a t a specified bmep. (Compare Curves 1680-1 and 1680-7 opposite .page 32.) They are commonly restricted to auto lean operation in the cruising range. A schematic example of a bmep curve is given above in Fig. 23. Referring to the schematic: on the altitude part of the curve solid horizontal lines, A, B, C, D, of constant bhp, rpm, and maximum bmep a t part throttle extend from sea level to the highest altitudes a t which the maximum bmep can be obtained with fulI throttle. These lines break a t the full throttle, maximum bmep line, EF, and become full throttle, constant rpm lines similar to thdse of the conventional altitude curve previously discussed. All operation within the area ADEF is at maximum bmep. Since maximum bmep is a

.... . -

4

1.

, 7

constant, the ratio of bhp/rpm is also a constant within this area - a fact borne out by inspection of the curve, which shows that the ratios 650 bhp/2200 rpm, 590 bhp/2000 rprn, ett., 'are approxiynateIy equal. The upper boundary line, AFG, is the maximum permissible bhp (650) for cruising. 'Below the altitude, F, a t which maximum bmep'is obtained, the maximum bhp line, AF, represents part throttle operation a t constant, rpmf above this altitude, F, the maximum bhp line, FG, represents full. throttle operation, with increasing .rpm and decreasing bmep. The, line terminates-'
TH.E U S E O F O P E R A T I N G C U R V E S

The following examples, based on the schematic above, will serve to illustrate the use of the bmep cruising curves: A pilot wishes to cruise a t maximum bmep and a bhp of 560. His maximum permissible cruising rprn is 2400.

1. To what rprn should he set his propeller governor? 2. Between what limits will he vary (by means of his throttle) the manifold pressure, and what value should the latter ' have at, say, 2000 ft.? 3. What is the greatest altitude attainable operating a t maximum bmep and 560 bhp? 4. If the pilot wishes to exceed this altitude and still operate a t 560 bhp, what must he do? What is the maximum altitude a t which he can cruise with 560 bhp? Solutions:

i -

-1.From the altitude curve it will be seen that the 560 bhp line lies about midway between the 1800 rprn and 2000 rprn part throttle, constant rprn and bhp, maximum bmep lines, C and B. Accordingly, a propeller governor setting of approximately 1900 rprn is indicated. 2. On the sea level curve, locate the intersection, P, of the given (560) bhp and required (1900) rpm. Its vertical projection indicates the manifold pressure a t sea level is 32.75 in. hg. On the altitude curve, locate the intersection, Q, of the 560 bhp - 1900 rprn line and the full throttle, maximum bmep line, EF. Both the solid and the dash manifold pressure lines indicate the manifold pressures a t Q is approximately 31 in. hg. 32.75 in. hg and 31 in. hg are therefore the limits within which the pilot must vary his manifold pressure by means of the throttle. The location of the intersection of the -560 bhp - 1900 rprn line and the 2000 ft. altitude line indicate that the manifold pressure a t 2000 ft. is approximately 32 in. hg. In this instance manifold pressure is shown by the solid lines. 3. Locate, as before, the intersection, Q, of the 560 bhp - 1900 rprn line and the full throttle, maximum bmep line, EF.

The vertical projection of this intersection shows that the greatest altitude attainable a t maximum bmep and 560 bhp is 4600 ft. 4. If the rprn (1900) is kept constant above this altitude (4600 ft.), the desired bhp (560) cannot be maintained, since the throttle can be opened no further. If 560 bhp is required above 4600 ft., it can be obtained only by gradually resetting the propeller governor to increase the rpm, a t full throttle, from 1900 rprn a t 4600 ft. to 2400 rpm, the maximum permissible cruising rpm, at 13500 ft., the critical altitude for 560 bhp and 2400 rpm. (Since bmep =KXbhp/rpm, it can be seer1 that the bmep decreases above the full throttle, maximum bmep line, EF. As shown by the line QR, the rpm increases while the bhp remains constant; accordingly the value of ratio bhp/rpm - and hence the bmep - decreases.) For accurate use, manifold pressures should be corrected for variations in c.a.t's. from standard. These corrections are made as before. Thus, in the example just given, had the c.a.t: been 10" C. (18"F.) below standard throughout, manifold pressures during cruise a t maximum bmep should have been reduced 0.5 in. - from 32.75 in. to 32.25 in. a t sea, level, and from-31in. to 30.5 in. at 4600 ft. Bmep Limits

The upper ends of the constant rprn lines on the sea level and altitude portions of many power curves of the conventional type are terminated by a heavy curved line. See, for example, Curve No. 1680-1 opposite page 32. Unless the contrary is indicated, these terminal lines are generally understood to be maximum permissible bmep lines, and are frequently designated as such, as on the curve referred to. From such a curve the operator can tell a t a glance the maximum manifold pressure which may be used under all conditions covered by the curve without exceeding bmep or bhp limits. This manifold pressure, however, does not necessarily give maximum performance under all conditions. Thus, referring again to Curve No. 1680-1, it will be seen that no operation a t 28 in. hg manifold pressure can result in excessive bmep's or bhp's so long as the rpm is below 2325, while the operations at 30.5 in.

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hg will similarly be within safe limits, except for a very small range above 700 bhp and 2150 rpm. (Curve No. 1680-1 is for auto lean operation, it must be noted, and the limits established by it do not apply to other types of operation, such as auto rich.) Excessive Ismep's cannot be obtained with an engine equipped with a fixed pitch propeller of proper design, since the load imposed by the propeller maintains a safe bhp/rpm ratio so long as the manifold pressure is kept within the maximum permissible limit. The flexibility introduced by the variable pitch propeller, however, may make it necessary to calculate the bmep for a certain set of conditions to see whether or not it

lies within the specified limits. For instance, referring once more to Curve No. 1680-1 opposite page 32, the operator may wish to know the bmep obtained a t 600 bhp and 1600 rpm. Substituting in the formula:

6-

792000 XBHP 792000 X600

BMEP

Displ. xRPM

=

1830 X1600

= 162 psi.

This is clearly in excess of the limiting cruising bmep of 140 psi, as specified on the curve - a fact which could have been obtained without computation by noting that the intersection of the 600 bhp and 1600 rpm lines lies within the restricted - prohibited area of both sea level and altitude curves.

PROBLEMS The problems which follow are for those who may desire further practice in the use of operating or power curves. The problems may be solved by using the typical curves, Nos. 1680-1, 1680-2, 1680-7, opposite page 32, as it will be assumed in all cases that the engine is a Twin Wasp R-1830SlC3G. The quantitative results obtained should agree withinf 2%y0 of the answers given on page 31. 1. (Curve No. 1680-2) An engine is run on a test stand a t 0 ft. altitude. The tachometer reads 2100 rpm; the manifold pressure gauge 23.25 in. hg; and the c.a.t. gauge 19" C. (66" F.) What is its bhp and bmep? 2. (Curve No. 1680-1) An airplane is cruising a t 11,500 ft., as shown by the altimeter (indexed to 29.92 in. hg). The tachometer and manifold pressure gauge read 1900 rpm and 26.5 in. hg, respectively. The c.a.t. is -23" C. ( - 10" F.) What is the actual bhp being developed by the engine? 3. (Curve No. 1680-2) Because of ram pressure an engine a t full throttle is able to , develop 2400 rpm and 34 in. hg manifold pressure a t 13,000 ft. Assuming standard c.a.t. conditions, what is the increase in critical altitude and bhp because of the ram? 4. (Curve No. 1680-1) A flight engineer wishes to cruise a t 6000 ft., 450 bhp, and 1850 rpm. C.a.t. is estimated to be -20" C. (-4" F.) At what point should the manifold pres-

5.

6.

7.

8.

9.

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sure be set? Will this be full throttle or part throttle operation? (Curve No. 1680-2) What manifold iressure will be required to fly a t 12,500 it., 2500 rpm, and 830 bhp, if c.a.t. is -26" C. ( - 15" F.)? (Curve No. 1680-7) A pilot wishes to cruise a t maximum bmep, 620 bhp, and 6000 ft. altitude. The c.a.t. at this altitude is known to be 14" C. (57" F.) What combination of rpm and manifold pressure should he select? What is the bmep? Is operation a t full or part throttle? (Curve No. 1680-7) If the pilot in the foregoing problem had wished to go to 13,100 ft. without increasing his bhp above 620, what rpm and manifold pressure would have been called for, assuming standard conditions? Would operation be a t full or part throttle? (Curve No. 1680-2) For an airplane to takeoff, its engine must develop 1100 bhp. The flying field is at 7000 ft. altitude; it is 2:00 PM; and the free air temperature (or c.a.t.) is 35" C. (95" F.) on the ground. Will'the airplane be able to take off, assuming these are the only determining factors? (Curve No. 1680-7) What is a safe maxi- . mum manifold pressure for all operations below 2230 rpm? Will maximum bmep ever be attained? (Assume standard conditions.)

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THE USE O F OPERATING CURVES

(2) What is the bmep of this combination? Is it within the max. continuous rating bmep limit? (3) Assuming standard conditions, what should the operator's manifold pressure be to deliver the specified 980 bhp at 2400 rprn at sea level? At the critical altitude? At 4000 ft.? (4) Had the c.a.t. been 7" C. (13" F.) above standard at each of these altitudes, what correction should be applied to the manifold pressure figures? c. .Would a climb a t 980 bhp and 2200 rprn be feasible? Explain.

10. (Curve No. 1680-1) Draw the part throttle, constant rpm, constant manifold pressure line for operation at 2000 rpm and 27 in. hg. manifold pressure. About half of this line will be seen to lie in the restricted-prohibited range. .Does this mean that operation at 2000 rprn and 27 in. hg. is forbidden below, say, 5000 ft.? (Assume standard conditions.) 11. (Curve No. 1680-2) a. What is the bmep a t maximum continuous rating - i.e. a t normal rated power and rpm? b. The operator wishes to climb using 980 bhp and 2400 rpm: (1) What is the critical altitude of this bhp-rpm combination?

ANSWERS 1. Bhp from sea level curve: 715 Actual bhp, corrected (-1%) for c.a.t.: 708. Answer 792000 708 X -= 146 psi. Answer 1830 2100 2. Bhp from altitude curve: 524 Actual bhp, corrected (+3%) for c.a.t.: 540. Answer 3. Increase in Critical Altitude: 13,000 ft. - 10,500 ft. =2500 ft. Answer Increase in bhp: 865 bhp - 856 bhp = 9 bhp. Answer 4. Manifold pressure from curve: 25.5 in. hg. Actual manifold pressure, corrected ( - 1in. hg.j for c.a.t.: 24.5 in. hg. Answer Part throttle operation. Answer 5. Manifold pressure from curve: 32.25 in. hg. Actual manifold pressure, corrected ( - 0.75 in. hg.) for c.a.t. : 31.5 in. hg. Answer 6. Rpm: 1915. Answer Manifold pressure from curve: 31.25 in. hg. Actual manifold pressure, corrected (f0.5 in. hg.) for c.a.t.: 31.75 in. hg. Answer Bmep: 140 bsi by computation; 140 psi from curve. Answer Part throttle operation. Answer 7. Rpm: 2100; Manifold Pressure: 28 in. hg. Answer Full throttle operation. Answer 8. No. Naximum available power for 5 min. a t 7000 ft. is 1115 bhp, under standard conditions, as shown by the curve, which Bmep =

would be enough for take-off. However, correcting this bhp for c.a.t., we find the actual figure is only 1048,which is inadequate for our purposes. Better wait till it cools off a bit. Answer 9. Manifold Pressure: 29 in. hg. Answer No. Answer 10. No. The maximum bhp obtainable under the given rpm-manifold combination is 575 (at the critical altitude). Accordingly the bmep's will always be below 140 psi. (actually below 125 psi). This fact is shown on the curve by the location of the intersection of the 27 in. hg manifold pressure line and the 2000 full throttle, constant rprn line, which intersection is to the right of the full throttle, maximum bmep line. Note also that the constant manifold pressure (27 in. hg) is below all limiting values. Answer 792000!x !?! ! =433 x 1050 -= 178 psi. Ans. 1830 2550 2550 b. (1) 6600 ft. Answer 980 = 177 psi. Answer (2) Bmep =433 X2400 Yes. I t is below max. continuous rating bmep (178 psi). Answer (3) 40.75 in. hg; 39.00 in. hg; 40.00 in. hg. Answer (4) Add 0.25 in. hg. Answer c. No. Bmep is 193, and is too great. Am.

PRATT & W H l T N E U AIRCRAFT

TYPICAL POWER CURVES

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(-swPERSEDES

1653)

CURVE NO. INST. 1680-1

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LIBRATION - HORSE POWER AND MANIFOLD PRESSURE AT STAP PHERIC CONDITIONS WITH AUTO- RI CH MiXTURE ,

Pratt & Whitney The Aircraft Engine and Its Operation_Rev1949_BZ

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