AIRCRAFT DESIGN PROJECT- I 7-15 SEATER BUSINESS JET
8
ABSTRACT In this project we have designed a 7-15 seater Business jet. We have taken the sufficient steps to make sure sure that that the the airc aircra raft ft what what we are are desi design gniing is in an opti optimu mum m rang range e. The The airc aircra rafft para parame mete ters rs like crui ruise velo veloci city, ty, cruis cruise e alti altitu tude de,, wing wing loadin loading g etc etc
and and weig weight ht
estimation, airfoil selection, wing selection, landing gear selection tion have been made with extre treme care are. The The adequate ate detail ails have ave been coll ollected to make our calc calcul ulati ation on easi easier er and and to make make desi design gn more more prec precis isio ion. n. The The detai details ls have have been been colle collect cted ed from from vario various us sourc sources es which are given in the bibliography.
8
ABSTRACT In this project we have designed a 7-15 seater Business jet. We have taken the sufficient steps to make sure sure that that the the airc aircra raft ft what what we are are desi design gniing is in an opti optimu mum m rang range e. The The airc aircra rafft para parame mete ters rs like crui ruise velo veloci city, ty, cruis cruise e alti altitu tude de,, wing wing loadin loading g etc etc
and and weig weight ht
estimation, airfoil selection, wing selection, landing gear selection tion have been made with extre treme care are. The The adequate ate detail ails have ave been coll ollected to make our calc calcul ulati ation on easi easier er and and to make make desi design gn more more prec precis isio ion. n. The The detai details ls have have been been colle collect cted ed from from vario various us sourc sources es which are given in the bibliography.
8
Serial no
Page no
Topic
1
Introduction
5
2
Classification of aircraft
9
3
The design
11
4
Comparative data sheet
12
5
Weight design
31
6
Wing design
35
7
Airfoil selection
37
8
Landing gear design
42
9
Performance calculations
53
10
3-D view Diagram
56
8
INTRODUCTION Purpose and scope of airplane design An airplane is designed to meet the functional, operational and safety requirements set by or acceptable to the ultimate user user.. The The actu actual al proce rocess ss of desi design gn is a comp comple lex x and and long long draw drawn n out out engineering task involving: •
Selection of airplane type and shape
•
Determination of geometric parameters
•
Selection of power plant
•
Structural design and analysis of various components and
•
Determination characteristics.
of
airplane
flight
and
operational
Over the year of this century, aircraft have evolved in many directions and the design of any modern plane is a joint project for a large large body body of comp compete etent nt engi enginee neers rs and and techn technici icians ans,, head headed ed by a chief chief designer. Different groups in the project specialize in the design of different components of the airplane, such as the wing, fuselage etc. A new experimental plane has to meet higher perform performance ance requirem requirements ents than similar planes planes already already in service. service. Hence Hence desi design gn labo labora rato tori ries es invo involv lved ed in expe experi rime ment ntal al and and rese resear arch ch work work are are indispensable indispensable adjuncts to a design office. These laboratories as well as allied specialized design offices and research institutions are concerned in helping the designer to obtain the best possible solutions for all problems pertaining to airpla airplane ne desig design n and and const constru ructi ction on and and in the the devel develop opme ment nt of suitab suitable le components and equipment. Airplane design procedure is basically a method of trial and erro errorr for the desig design n of comp compon onen entt units units and and their their harmo harmoni nizat zatio ion n into into a complete aircraft system. Thus each trial aims at a closer approach to the
8
final goal and is based on a more profound study of the various problems involved. The three phases of aircraft design are •
Conceptual design
•
Preliminary design
•
detail
Phase of aircraft design Conceptual design Aircraft design can be broken into three major phases, as depicted in figure. Conceptual design is the primary focus of this book. It is in conceptual design that that the the basi basic c ques questi tion ons s of confi configu gura rati tion on arra arrang ngem emen ent, t, size size and and weig weight ht,, and and performance are answered. The first question is “can an affordable aircraft be built that meets the requirements?” if not, the customer may may wish to relax the requirements. Conceptual design is a very fluid process. New ideas and problems emerge as a design is investigated in increasing detail. Each time the latest design is analyzed and sized, it must be redrawn to reflect the new gross weight, fuel weight, wing size, and other changes. Early wind tunnel test often revel problems requiring some changes to the configuration.
REQUIREMENTS
CONCEPTUAL DESIGN
will it work? What does it look like? What requirements drive the design? What trade-offs should be consider? What should it weigh and cost?
8
PRELIMINARY DESIGN
freeze the configuration Develop lofting Develop test and analytical base Design major items Develop actual cost estimation
DETAIL DESIGN
design the actual piece
to be built Design the tooling and fabrication process Test major items structure, landing gear Finalize weight and performance estimate
FABRICATION
Preliminary design
Preliminary design can be said to begin when the major changes are over. The big questions such as whether to use a canard or an aft tail have been resolved. The configuration arrangement can be expected to remain about as shown on current drawing, although minor revisions may occur. At some point late in preliminary design, even minor changes are stopped when a decision is made to freeze the configuration. During preliminary design the specialists in area such as structure landing gear and control systems will design and analyze their portion of the aircraft. Testing is initiated in areas such as aerodynamics, propulsion, structures, and control. A mockup may be constructed at this point. A key activity during preliminary design is “lofting”. Lifting is the mathematical modeling of the outside skin of the aircraft with sufficient accuracy to insure proper fit between its different parts, even if they are designed by different designers and possibly fabricated in different location. Lofting originated in shipyards and was originally done with long flexible rulers called” splines”. This work was done in a loft over the shipyard; hence the name.
8
The ultimate objective during preliminary design is to ready the company for the detail design stage, also called full-scale development. Thus, the end of preliminary design usually involves a full scale development proposal. In today’s environment, this can result in a situation jokingly referred to as “you-bet-your-company”. The possible loss on an overrun contrast o from lack of sales can exceed the net worth of the company! Preliminary design must establish confidence that the airplane can be built in time and at the estimated cost. Detailed design
Assuming a favorable decision for entering full scale development, the detail deign phase begins in which the actual pieces to be fabricated are designed. For example, during conceptual and preliminary design the wing box will be designed and analyzed as a whole. During detail design, that whole will be broken down in to individual ribs, spars and skins, each of which must be separately designed and analyzed. Another important part of detailed is called production design. Specialist determine how the airplane will be fabricated, starting with the smallest and simplest subassemblies and building up to the final assembly process. Production designers frequently wish to modify the design for ease of manufacture; that can have a major impact on performance or weight. Compromises are inevitable, but the design must still meet the original requirements. It is interesting to note that in the Soviet Union, the production design is done by a completely different design bureau than the conceptual and preliminary design, resulting in superior producibility at some expense in performance and weight. During detail design, the testing effort intensifies. Actual structure of the aircraft is fabricated and tested. Control laws for the flight control system arte tested on an “iron-bird” simulator, a detailed working model of the actuator and flight control surfaces. Flight simulator are developed and flown by both company and customer test pilot. Detail design ends with fabrication of the aircraft. Frequently the fabrication Begins on part of the aircraft before the entire detail-design effort is completed. Hopefully, changes to already- fabricated pieces can be avoided. The further along a design progresses, the more people are
8
involved. In fact, most of the engineers who go to work for a major aerospace company will work in preliminary on detail design.
Classification of airplanes design Functional classification: The airplane today is used for a multitude of activities in civil and military fields. Civil applications include cargo transport, passenger travel, mail distribution, and specialized uses like agricultural, ambulance and executive flying. The main types of military airplane at the present time are fighters and bombers. Each of these types may be further divided into various groups, such as strategic fighters, interceptors, escort fighters, tactical bombers and strategic bombers. There are also special aircraft, such as ground attack planes and photo-re-connaisance planes. Sometimes more than one function may be combines so that we have multipurpose airplanes like fighter-bombers. In addition to these, we have airplanes for training and sport.
Classification by power plants: Types of engines used for power plant: •
Piston engines (krishak, Dakota, super constellation)
•
Turbo-prop engines ( viscount,friendship,An-102)
•
Turbo-fan engines (HJT – 16, Boeing series, MIG-21)
•
Ramjet engines
•
Rockets (liquid and solid propellants) (X-15A)
Location of power plant: •
Engine ( with propeller) located in fuselage nose (single engine) (HT-2,Yak-9,A-109)
•
Pusher engine located in the rear fuselage (Bede XBD2)
•
Jet engines submerged in the wing
8
1. At the root(DH Comet, Tu-104,Tu-16) 2. Along the span (Canberra, U-2, YF-12A)
•
Jet engines in nacelles suspended under the wing (pod mountings) (Boeing 707,DC-8,Convair 880)
•
Jet engines located on the rear fuselage (Trident, VC – 10 ,i1-62)
•
Jet engines located within the rear fuselage (Hf – 24, lighting,MIG-19)
Classification by configuration:
Airplanes are also classified in accordance with their shape and structural layout, which in turn contribute to their aerodynamic, tactical and operational characteristics. Classification by configuration is made according to: • • •
Shape and position of the wing Type of fuselage Location of horizontal tail surfaces
Shape and position of the wing: •
Braved biplane(D.H. Tiger moth)
•
Braced sesquiplane (An-2)
•
Semi-cantilever parasol monoplane (baby ace)
•
Cantilever low wing monoplane (DC-3,HJT-16,I1-18,DH Comet)
•
Cantilever mid wing monoplane (Hunter, Canberra)
8
•
Cantilever high wing monoplane (An-22,Brequet 941 Fokker Friendship)
•
Straight wing monoplane (F-104 A)
•
Swept wing monoplane (HF-24, MIG-21, Lighting)
•
Delta monoplane with small aspect ratio (Avro-707, B-58 Hustler, Avro Vulcan)
Type of fuselage • • •
Conventional single fuselage design ( HT-2,Boeing 707 Twin- fuselage design Pod and boom construction (Packet, Vampire)
Types of landing gear: •
Retractable landing gear (DC-9,Tu-114,SAAB-35)
•
Non- retractable landing gear (pushpak, An-14, Fuji KM-2)
•
Tail wheel landing gear (HT-2,Dakota,Cessana J85 C)
•
Nose wheel landing gear (Avro-748, Tu-134,F-5A)
•
Bicycle landing gear (Yak-25,HS-P,112)
THE DESIGN Design is a process of usage of creativity with the knowledge of science where we try to get the most of the best things available and to overcome the pitfalls the previous design has. It is an iterative process to idealism toward with everyone is marching still. Design of any system is of successful application of fundamentals of physics. Thus the airplane design incorporates the fundamentals of aerodynamics, structures, performance and stability & control and basic physics. These are based on certain degree of judgment and experience. Every designer has the same technical details but each design prevails it own individuality and the mode of the designer.
8
Here the preliminary design has been done of an executive Transport Aircraft. The basic requirements are the safe, comfortable and economic transport mode with reasonable time period of flight. Here comfort and safety are given primary importance. Here the most possible considerations have been taken. And the flight parameters and limitations are studied.
The modern day calls for the need of latest aircrafts for the use of passenger transport which aims mainly at improving the aerodynamic characteristics as well as the passenger comfort. This design project also looks at the above aspects in a lot more closer way. Also the design project has been classified into different stages in our design will be as follows.
•
Collection of comparative data
•
Selection of aircraft parameters
•
Preliminary and second weight estimations
•
Selection of power plant
•
Airfoil selection, flaps, t/c, sweep, etc
•
Layout of L/G, load, tires selection
•
3-view diagram
•
Balance diagram
8
COMPARATIVE DATA SHEET
In the designers perspective it is necessary to compare the existing airplanes that are of same type as that of our desired airplane. Their important parameters, positive aspects to be considered and pitfalls to be overcome are taken into consideration.
The datas has been collected for various sites in the internet for 7-15 seater BUSINESS JET design Several parameters are compared for over 25 aircrafts and different critical parameters were plotted on graph. They are •
Cruise Velocity
•
Range
•
Wing area
•
Wing loading
•
Aspect ratio
•
Sweep angle
•
Cruise altitude
• •
Thrust/weight ratio Maximum take off weight
8
Parameters
Bombardier global express
Britten-Norman Cessna 421 Islander
Maximum take off Weight (kg)
44,500
2,994
3,103
Empty weight(kg)
22,600
1,627
2,132
Design pay load(kg)
805
n/a
n/a
Design fuel load(kg)
19663
n/a
n/a
Length(m)
30.3
10.86
10.29
Area(m2)
94.9
30.2
18.6
28.65
14.94
12.15
8.65
7.39
7.937
Span(m) Aspect ratio Quarter chord Sweep(degrees) Max wing loading(kg\m 2) Thrust\weight ratio Length\span ratio Cruise velocity(km\hr) Design range(km) Altitude(m)
Capacity
35
n/a
n/a
468
99.14
166.8
0.301
0.186
0.158
1.057
0.7269
0.847
950
n/a
420
11,390
1,400
2,756
15,500
4,024
5,230
8 to 19
9
8
8
Parameters Maximum take off Weight(kg)
Cessna citation excel
Cessna citation cj3
6,291
Cessna citation X
13,608
16,375
Design pay load(kg)
371
Design fuel load(kg)
2,136
3057
Length(m)
15.29
19.37
22
27
34.35
50
16.26
19.24
19.4
10
8.4
7.527
0
0
37
230
264.1
334.45
0.331
0.38
0.33
0.94
1.006
1.13
Cruise velocity(km\hr)
773
848
1004
Design range(km)
3519
4,821
6,020
13,716
15,517
15,545
9
10
Area(m2) Span(m) Aspect ratio Quarter chord Sweep(degrees) Max wing loading(kg\m 2) Thrust\weight ratio Length\span ratio
Altitude(m)
Capacity
6,486
n/a
8
n/a
8 to 12
Parameters
Gulf stream G100
Maximum take off Weight(kg)
Gulf stream G200
Hawker 400
11181
16079
7393
6214
8981
4785
Design pay load(kg)
363
1905
289
Design fuel load(kg)
3942
6804
2228
Length(m)
16.94
18.97
14.75
Area(m2)
29.41
34.28
22.43
Span(m)
16.64
17.7
13.25
8.8
9.1
7.8
34
34
20
380.3
469.1
329.67
0.31
0.346
0.287
1.01
1.07
1.11
850
880
778
5461
6667
2883
12497
13715
13243
9
8
8
Empty weight(kg)
Aspect ratio Quarter chord Sweep(degrees) Max wing loading(kg\m 2) Thrust\weight ratio Length\span ratio Cruise velocity(km\hr) Design range(km) Altitude(m)
Capacity
Parameters
Hawker 800
Dassault Falcon 50
8
Learjet 25
Maximum take off Weight(kg)
12701
18007
Empty weight(kg)
7029
9603
Design pay load(kg)
1000
1710
n/a
Design fuel load(kg)
4536
7040
n/a
15.6
18.52
14.5
Area(m2)
34.75
46.83
21.53
Span(m)
15.66
18.86
10.84
7.1
7.6
5.4577
20
29
365.5
384.5
Length(m)
Aspect ratio Quarter chord Sweep(degrees) Max wing loading(kg\m 2) Thrust\weight ratio
n/a
Length\span ratio Cruise velocity(km\hr) Design range(km) Altitude(m)
Capacity
Parameters Maximum take off Weight(kg)
n/a
6803
n/a
315.9
n/a
0.99
0.98
793
850
4704
5602
2853
11887
12497
13720
14
9
8
Cessna 680 citation sovereign
Learjet 45 9298
13067
8
1.3376
n/a
Cessna 550 citation 6713
Empty weight(kg)
5797
7892
3992
Design pay load(kg)
1030
726
370
Design fuel load(kg)
2750
4921
2204
Length(m)
17.81
19.35
14.4
Area(m2)
28.95
49.73
30
Span(m)
14.56
19.25
15.75
7.3
7.7
8.4
13
12
10
Max wing loading(kg\m 2)
322.3
283.9
223.8
Thrust\weight ratio
0.383
0.33
0.38
1.22
1.005
0.9
846
850
744
3885
5056
1744
15545
13105
8458
9
12
10
Aspect ratio Quarter chord Sweep(degrees)
Length\span ratio Cruise velocity(km\hr) Design range(km) Altitude(m)
Capacity
Parameters
Learjet 35
IAI WEST WIND
Maximum take off Weight(kg)
8300
Empty weight(kg)
4590
GAF N22 NOMAD 9390
n/a
8
3855
n/a
Design pay load(kg)
n/a
n/a
n/a
Design fuel load(kg)
n/a
n/a
n/a
Length(m) Area(m2) Span(m) Aspect ratio Quarter chord Sweep(degrees)
12.56
23.5
28.6
30.1
12.04
13.65
16.52
6.16
6.5
9
n/a
Thrust\weight ratio Length\span ratio Cruise velocity(km\hr) Design range(km) Altitude(m)
n/a
353.19
328.322
128.07
0.383
0.3
0.192
1.23
1.16
0.76
852
870
311
4070
3410
1352
13716
13716
6400
10
12
8
Parameters
RAYTHON BEECHCRAFT KING AIR
Maximum take off Weight(kg) Empty weight(kg)
15.93
n/a
Max wing loading(kg\m 2)
Capacity
14.83
Cessna citation mustang
Learjet 23
6350
n/a
8
6803
3807.97
n/a
n/a
Design pay load(kg)
n/a
n/a
281
Design fuel load(kg)
n/a
n/a
1460
Length(m) Area(m2) Span(m) Aspect ratio Quarter chord Sweep(degrees)
28.2
21.46
22.3
16.61
13.18
13.16
9.78
8.09
7.7
n/a
n/a
0.8
Cruise velocity(km\hr)
215.6
n/a
0.82246
583
Design range(km)
n/a
317
0.155
Length\span ratio
Parameters
13.16
225.18
Thrust\weight ratio
Capacity
10.84
n/a
Max wing loading(kg\m 2)
Altitude(m)
13.36
1
n/a
3630
N\A
630
1584
2310
13715
12500
6
4 to 5
15
Piaggio P180 Avanti
North American sabreliner
Lockheed jet star
Maximum take off Weight(kg)
8060
5262
20227
Empty weight(kg)
5000
3400
11477
Design pay load(kg)
n/a
n/a
8
n/a
Design fuel load(kg)
n/a
1271
n/a
Length(m)
13.41
14.4
18.42
Area(m2)
31.78
16
50.4
Span(m)
13.56
14.03
16.6
5.7858
12.3
5.46
Aspect ratio Quarter chord Sweep(degrees)
n/a
1
n/a
Max wing loading(kg\m 2)
253.6
327.4
401.3
Thrust\weight ratio
0.338
0.333
0.339
0.9889
1.026
1.109
644
823
2595
4585
12500
13135
9
8to 10
Length\span ratio Cruise velocity(km\hr)
n/a
Design range(km) Altitude(m)
Capacity
4020
n/a
4 to 5
8
Parameters
Learjet 60
Maximum take off Weight(kg)
Dassault falcon 7X
Learjet 55 10660
9525
28,894
Empty weight(kg)
6641
5832
n/a
Design pay load(kg)
1070
n/a
Design fuel load(kg)
3588
n/a
n/a
17735
17.88
16.8
23.19
24.6
24.57
70.7
Span(m)
13.34
13.35
25.13
Aspect ratio
7.234
7.253
8.93
Length(m) Area(m2)
Quarter chord Sweep(degrees) Max wing loading(kg\m 2) Thrust\weight ratio Length\span ratio Cruise velocity(km\hr) Design range(km) Altitude(m)
Capacity
n/a
34
n/a
433
387.66
435
0.392
0.345
0.33
1.34
1.258
0.923
782
744
685
4628
4156
10,500
15545
15545
15,500
8 to 10
7
19
8
8
8
8
8
s.n Design o characteristics
Values
1
Cruise velocity
800 Km/hr
2
Design range
3825 Km
3
Cruise altitude
12750ft
4
Aspect ratio
7.5
5
Sweep angle
31
6
Thrust ratio
.33
7
Wing loading
8
Max take weight
9
Wing area
/weight
315 Kg/M2 off
11250 kg
30
8
Weight Estimation The weight of the aircraft (W) is the key factor in almost aircraft performance problems. The gross weight is distributed in the following manner: W = Wstruc + Wcrew + Wpass + Wfe + Wpp + Wf
Here, Wstructure consists of the wing, fuselage, under-carriage & the empennage and accounts for about 32% of the gross weight, i.e., 0.32W. Wfixed equipment includes the passenger seats, food, baggage racks, lavatories, air-conditioning, avionics and other passenger amenities. This adds to the weight by about 0.05W. Wpowerplant is the weight of the engine and its systems. The initial assumption of engine weight is assumed to be 0.055W which may be modified later to suit thrust requirements. Wfuel is the weight contribution of the fuel to the total weight. It depends on the range also includes the Reserve fuel that is used in case of an emergency. It adds to the gross weight by a factor of 0.3W. Wcrew + Wpassengers accounts for the remaining weight. i.e., 0.275W. Taking passenger & baggage weight into consideration, a maximum of 1800N per passenger is permissible. As for a crew member, 1000N would suffice.
•
W = 0.725W (Crew)*1000N
+
(No.
8
of
passengers)*1800N
+
•
As for the aircraft to be designed, the Number of passengers will be 15 with a 5 member crew.
•
W = 116363 N = 11861.7 kg (approx)
•
The total weight of the power-plant (0.055W) requires being approx. 326 kg.
• The choice of a suitable engine, having been made, it is now
possible to estimate the amount of fuel required for a flight at the given cruising speed for the given range. •
Wfuel = (no. of engines) x (thrust at altitude) x Range x SFC x 1.2 ------------------------------------------
---------------Cruise velocity •
The factor of 1.2 is provided for reserve fuel
• Thrust at altitude is calculated using the relation
T σ
=
T 0 * σ
1.2
σ
ρ alt
=
ρ 0
• Therefore the weight of the fuel
Wfuel = 3385.43 kg
Final weight estimation:
W = 0.425W + Wpayload +Wfuel
8
W = 0.425W + 3261.97 + 3385.43
Hence W = 10821.57 kg
This value closely matches with maximum take off weight obtained in the comparative data sheet obtained. Engine selection: Therefore the weight of the power plant = 652 kg Choice of engine is a Turbofan for obvious reasons such as higher operating fuel economy & efficiency for high payloads. Since it is a business jet of 7 to 15 seater hence two engines would be sufficient. Engines can be used in combination of 2 x 326 kg engines ENGINE NAME
DRY WEIGHT (Kg)
SFC (lb/hr/lbt)
MAX. THRUST(KN)
TFE731-5A
400
0.469
20
J60-P-6
203
0.96
13.3
PW530A
275
0.468
18
CF700-202
349
0.65
20
JT15D-5
291
0.54
12.9
J85-GE-21
310
1.24
15.6
The power plant chosen is CF700-202 Wstruc = 0.32 Wg = 3600 kg Wfixed equipment = 0.05 Wg = 562 kg
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Wpower plant = 0.055 Wg = 652 kg Wfuel = 3385.43 kg Wcrew + Wpassengers = 3261.9 kg Wtotal = 11460 kg From data sheet we know that T/W = 0.33 Hence thrust required = 37 K Parameters :MTOW (N)
112435.6
Wpay (N)
32000
Wst (N)
35316
Wfe (N)
5513
Wpowerplant (N)
6396
Wfuel (N)
33211
Required Thrust (kN)
37
Thrust per engine (kN)
20
Cruise altitude (m)
11750
Sigma
0.26
Range (km)
3825
Cruise Velocity (kmph)
800
Time of cruise (hrs)
10
Power plant chosen
CF700-202
Dry weight (kg)
349
Net engine weight (N)
3423.7
Thrust per engine (kN)
20
Net Thrust (kN)
40
8
Cruise SFC (lb/hr/lbt)
0.65
Calculated W fuel (N)
3375
Wing selection Λ can be approx reasonably with more simply constructed trapezoidal using with a taper ratio of
λ=0.5(assuming) S=30m2 A= (b2/S) A=7.5= (b2/S) b=
= 15m
Formula to calculate root chord is given by Cr=2b/ (A (1+λ)) =2
/ (7.5(1+0.5)) =2.67m
Ct/Cr=0.31
Ct=2.67 0.5=1.33m
The mean aerodynamic chord length is given C= 2Cr (1+λ+λ2)/ (3(1+λ)) =2 x 2.67(1+0.5+0.5 2)/ (3 1.5) C=2.08m The normalized span wise location of the mean aerodynamic chord from the centre span of the wing is (Y/b)= (1+2λ)/ (6 (1+ λ))
8
(Y/b)= (1+1)/ (6 1.5) =0.222m Mach number for the a/c is given by Mcruise =800/1188=0.673 McrD = Mcruise +0.071 = 0.744 McrDo = McrD + ΔMcrD (AR) - ΔMcrD (CL) = 0.748 From the historical data for M, ΛLE=31⁰ Λc/4=tan-1(tan 31-((2 Cr (1- λ))/4b) Cr =2.67m b=15m λ=0.5 Λc/4=29.5⁰ Aspect Ratio(AR)
7.5
Area(S)
30M2
Span(b)
15M
Cr
2.67M
Ct
1.33M
λ
0.5
Λc/4
29.5⁰
AIRFOIL
SELECTION
During steady level flight, lift is equal to weight and hence
8
L = W = 0.5 ×density × V2stall × S×CLmax Density = 1.2256 kg/m 3 Vstall = 1.25 × V cruise Vcruise = 222.22 m/s (from graph) Vstall = 0.25×222.22 = 55.555 m/s S = 30 m 2 (from graph) W = 11460kg Therefore 11460 ×9.81 = 0.5×1.2256×55.555 2 ×30× CLmax CLmax = 1.98
Airfoil selection:
From the aerofoil data book various airfoils of required t/c are taken and are tabulated for maximum lift coefficient and minimum drag. From the table the airfoil with optimum combination of maximum lift coefficient and minimum drag coefficient is selected. Airfoil
CLmax
CDomin
L/D
652015
1.4
0.004
350
652215
1.5
0.004
375
652415
1.6
0.004
400
0.004
400
652415 0.5)
(a
= 1.6
Airfoil selected (root): NACA 652415 8
All airfoils listed below are from NACA series Airfoil
CLmax
CDomin
L/D
63210
1.55
0.0045
344
63A010
1.2
0.0045
226
63A210
1.425
0.00425
335
64210
1.45
0.004
362.5
64110
1.4
0.004
350
64A010
1.225
0.00425
288
64A210
1.425
0.004
356.27
64A410
1.625
0.00425
382
65210
1.4
0.00375
373
65410
1.525
0.00375
406
66210
1.2
0.003
400
Selected airfoil (mean): NACA 64A410 6% Airfoil
CLmax
CDomin
L/D
63006
0.9
0.004
225
64006 64206 65006 65206
0.86 1.1 0.95 1.1
0.00375 0.0038 0.0034 0.0035
229.3333
Airfoil selected (tip): NACA 65206
Cr = 2.666m
8
=
2 S b(1 +λ )
289.4737 279.4118
314.2857
Ct
= 0.5 × C r = 1.333 m
Cm
=
C r
+ C t
2
=2m
parameter
root
mean
Tip
Chord (m)
2.666
2.0
1.333
NACA
652415
64A410
65206
t/c
15 %
10 %
6%
Clmax
1.6
1.625
1.1
Cdomin
0.004
0.00425
0.0035
Volume of the fuel that can be stored in the wing
V = (0.5×Cm × t/c × Cm ×b/2 × 0.75) ×0.75×2 = 1.68m3 So volume of fuel that can be stored in wings is 1.68 m3 Total Volume of the fuel to be carried Vt =
W fuel (9.81 * ρ fuel )
= 4.21 m3 Remaining Fuel
=
4.218 –1.68 = 2.58 m 3
The remaining fuel is stored in the fuselage.
8
Flap selection:
With zero final velocity and with deceleration aided by thrust reverser, landing speed is calculated using the following equation. Since our aircraft is a 15 seater aircraft, Runway length = 3100 m Ground run= 60% of (Runway length) = 1850m V2-VL2=2 a S V=0, a= -0.25g s=1850m VL = 95.25 m/s. Vstall = VL /1.3= 73.27m/s From the stalling speed the lift coefficient required to avoid stall is calculated as follows, CL stall =
2(W / S ) ρ Vstall
2
= 0.95
CLmaxavailable = 0.95 CLrequired ∆
CLreq
= 1.98 = 1.03
Part span correction required for flap:
The flap on the wing of the aircraft does not run along the entire span. As a convention we can approximately take the flap is provided for quarter span of the wing i.e. flap span is b/4. Hence ∆
∆
CLactual with the part span correction is given by,
CLactual =∆ CLrequired /4 (since the flap is of quarter span)
8
Length Cf
= 50% of Semi span = 0.25 C
Split Flap: Flap deflection
Full span CL
Part span 50%
CDo
CL
CDo
0
0
0
0
0
10
0.22
0.018
0.11
0.009
20
0.42
0.048
0.21
0.024
30
0.6
0.09
0.3
0.045
40
0.65
0.136
0.325
0.068
50
0.79
0.18
0.395
0.09
60
0.825
0.22
0.4125
0.11
LANDING GEAR SELECTION In aviation, the undercarriage or landing gear is the structure (usually wheels) that supports an aircraft and allows it to move across the surface of the Earth when it is not flying. Overview
Landing gear usually includes wheels equipped with shock absorbers for solid ground, but some aircraft are equipped with skis for snow or floats for water, and/or skids or pontoons (helicopters). Types of gear arrangements
Wheeled undercarriages come in two types: conventional or "tail dragger" undercarriage, where there are two
8
main wheels towards the front of the aircraft and a single, much smaller, wheel or skid at the rear; or tricycle undercarriage where there are two main wheels (or wheel assemblies) under the wings and a third smaller wheel in the nose. Most modern aircraft have tricycle undercarriages. Tail draggers are considered harder to land and take off, and usually require special pilot training. Sometimes a small tail wheel or skid is added to aircraft with tricycle undercarriage, in case of tail strikes during take-off. The Concorde, for instance, had a retractable tail “bumper” wheel (as delta winged aircraft need a high angle when taking off). Some aircraft with retractable conventional landing gear have a fixed tail wheel, which generate minimal drag and even improve yaw stability in some cases. Retractable gear
To decrease drag in flight some undercarriages retract into the wings and/or fuselage with wheels flush against the surface or concealed behind doors; this is called retractable gear. A design for retractable landing gear was first seen in 1876 in plans for an amphibious monoplane designed by Frenchmen Alphonse Pénaud and Paul Gauchot. Aircraft with at least partially retractable landing gear did not appear until 1917, and it was not until the late 1920s and early 1930s that such aircraft became common. By then, aircraft performance was improved to the point where the aerodynamic advantage of a retractable undercarriage justified the added complexity and weight. An alternate method of reducing the aerodynamic penalty imposed by fixed undercarriage is to attach aerodynamic fairings (often called "spats" or "pants") on the undercarriage, with only the bottoms of the wheels exposed.
Large aircraft
As aircraft grow larger, they employ more wheels to cope with the increasing weights. The Airbus A340-500/-600 has an 8
additional four-wheel undercarriage bogie on the fuselage centerline. The Boeing 747 has five sets of wheels, a nose-wheel assembly and four sets of four-wheel bogies. A set is located under each wing, and two inner sets are located in the fuselage, a little rearward of the outer bogies. Unusual types of gear
Some planes use wheels only for take off and drop them afterwards to gain the improved streamlining without the complexity, weight and space requirements of a retraction mechanism. In this case, landing is achieved on skids or similar simple devices. Historical examples include the Messerschmitt Me 163 and the Messerschmitt Me 321. A related contemporary example are the wingtip support wheels ("Pogos") on the U-2 reconnaissance aircraft, which fall away after take-off; the aircraft then relies on titanium skids on the wingtips for landing. Landing gear on an Airbus A310 an unusual undercarriage configuration is found on the Hawker Siddeley Harrier, which has two main wheels in line astern under the fuselage (called a bicycle or tandem layout) and a smaller wheel near the tip of each wing. On second generation Harriers, the wing is extended past the outrigger wheels to allow greater war loads to be carried. A multiple tandem layout was used on some military jet aircraft during the 1950s such as the Lockheed U-2, Myasishchev M-4, Yakovlev Yak-25, Yak-28 and the Boeing B-47 because it allows room for a large internal bay between the main wheels. A variation of the multi tandem layout is also used on the B-52 Stratofortress which has four main wheel bogies underneath the fuselage and a small outrigger wheel supporting each wing-tip. The B-52's landing gear is also unique in that all four pairs of main wheels can be steered. This allows the landing gear to line up with the runway and thus makes crosswind landings easier (using a technique called crab landing).For light airplanes, a landing gear which is economical to produce is a simple wooden arch laminated from ash, as used on some homebuilt aircraft. A recent
8
addition to this type of gear is the fixed-gear RJ.03 IBIS canard homebuilt aircraft. Steering
The steering mechanism used on the ground with wheeled landing gear varies by aircraft, but there are several general types of steering. Tail dragger aircraft may be steered by rudder alone (depending upon the prop wash produced by the aircraft to turn it) with a freely-pivoting tail wheel, or by a steering linkage with the tail wheel, or by differential braking (the use of independent brakes on opposite sides of the aircraft to turn the aircraft by slowing one side more sharply than the other). Aircraft with tricycle landing gear usually have a steering linkage with the nose wheel (especially in large aircraft), but some allow the nose wheel to pivot freely and use differential braking and/or the rudder to steer the aircraft. Virgin Atlantic Airbus A340-600 landing. This airliner has an undercarriage on the fuselage belly, as well as on the wings. Some aircraft require that the pilot steer by using rudder pedals; others allow steering with the yoke or control stick. Some allow both. Still others have a separate control, called a tiller , used for steering on the ground exclusively.
Rudder steering
When an aircraft is steered on the ground exclusively using the rudder, turning the plane requires that a substantial airflow be moving past the rudder, which can be generated either by the forward motion of the aircraft or by thrust provided by the engines. Rudder steering requires considerable practice to use effectively. Although it requires air movement, it has the
8
advantage of being independent of the landing gear, which makes it useful for aircraft equipped with fixed floats or skis. Direct steering
Some aircraft link the yoke, control stick, or rudder directly to the wheel used for steering. Manipulating these controls turns the steering wheel (the nose wheel for tricycle landing gear, and the tail wheel for tail draggers). The connection may be a firm one in which any movement of the controls turns the steering wheel (and vice versa), or it may be a soft one in which a spring-like mechanism twists the steering wheel but does not force it to turn. The former provide positive steering but make it easier to skid the steering wheel; the latter provide softer steering (making it easy to over control) but reduce the probability of skidding the wheel used for steering. Aircraft with retractable gear may disable the steering mechanism wholly or partially when the gear is retracted. Differential braking
Differential braking depends on asymmetric application of the brakes on the main gear wheels to turn the aircraft. For this, the aircraft must be equipped with separate controls for the right and left brakes (usually on the rudder pedals). The nose or tail wheel usually is not equipped with brakes. Differential braking requires considerable skill. In aircraft with several methods of steering that include differential braking, differential braking may be avoided because of the wear it puts on the braking mechanisms. Differential braking has the advantage of being largely independent of any movement or skidding of the nose or tail wheel. Tiller steering
A tiller in an aircraft is a small wheel or lever, sometimes accessible to one pilot and sometimes duplicated for both pilots, that controls the steering of the aircraft while it is on the ground. The tiller may or may not be designed to work in combination 8
with other controls such as the rudder or yoke. In large airliners, for example, the tiller is often used as the sole means of steering during taxi, and then the rudder is used to steer during take-off and landing, so that both aerodynamic control surfaces and the landing gear can be controlled simultaneously when the aircraft is moving at aerodynamic rates of speed. Maximum take off weight (from data sheet) = 10821.57 N= 25165 lbs Tricycle-wheel arrangement Nose- 1L/G - 2 wheels
Main – 2 L/G - 8 wheels
Weight taken by main landing gear, Wm = 0.9 ×W/8 = 12.6 KN/wheel = 2831 lbs/wheel Weight taken by nose landing gear, Wn = 0.1 × W/2 = 5600 N/wheel = 1258.25 lbs/wheel Tyre sizing:
During landing and takeoff, the undercarriage supports the total weight of the airplane. Undercarriage is of three types: 1. Bicycle type 2. Tricycle type 3. Tricycle tail wheel type. A tricycle wheel type needs more takeoff distance and floor is also needs to be inclined. So we have selected a tricycle nose wheel type. Also the types of runways are also be selected with
8
due care since depending on this criterion, wheels and tires are selected. Being an executive transport vehicle, it may land on grass fields and may even on beaches. For tricycle nose wheel type undercarriage, the nose wheel carries 10 % of the total load and the main undercarriage carries 90% of the total load. For different runways, the allowable loadings are given by, Grass
21.1 N/cm2
Grass Strip
36.9 N/cm2
Asphalt (Tar)
73.9 N/cm2
Concrete
116.1 N/cm2
Wheel configurations: Wheel diameter Wheel width
A 1.51 0.715
B 0.349 0.312
Main wheels:
Wheel diameter = A (Wm\2)
B
= 1.51× (2831\2)
0.349
= 19 inches = 0.48 meters Wheel width = A (Wm\2) B =0.715× (2831\2)
0.312
= 6.87 inches = 0.1746 meters Nose wheels:
Wheel diameter = A (Wn)B = 1.51× (1258.25) 0.349 = 18.23 inches = 0.463 meters
8
Wheel width = A (Wn)B = 0.715 × (1258.25) 0.312 = 6.628 inches = 0.168 meters. FUSELAGE SELECTION
Length of the fuselage = 16 m Diameter of the fuselage = 1.74 m HORIZONTAL TAILPLANE
From the standard design books =0.5 ARht= 4 Sht/s=0.15 S=30m2 Sht=4.5m2 bht= bht=4.24m Crt= (2Sht)/ ( +1) bnt)
= (2 4.5)/(1.5 4.24)
Crt=1.415m Ctt=0.5Crt=0.7m Yht= (bt (1
))/(6(1+ )) =(4.24(1+2(0.5))/(6(1+0.5))=0.94m
Cht= ((2Crt (1+ + 2))/ (3(1+ ) = (2 1.415(1+0.5+0.52)/ (3 (1+0.5))Cht=1.1m
Vertical tail plane
From standard design books Svt/S=0.08 Svt=2.4m bvt=
8
ARvt=1.5
bvt=1.879m Crt= (2 Svt / ((1+ ) bvt)) = ((2
/(1+0.5) 1.897))
Ctt=Crt0.5=0.84m Yvt = (2 bvt (1
))/ (6(1+ )) =0.843m
Cvt= ((2Crt (1+ + 2))/ (3(1+ ) =1.3m
Horizontal Tail:
VHT
=
ARht
0.7 =
4
SHT
=
4.5 m2
bht
=
4.24 m
Cr
=
1.415 m
Cmht = Ct
1.1 m =
Yht
0.7 m =
0.94 m
Vertical Tail:
VVT
=
ARvt SVT
0.04 =
=
1.5 2.4 m2
bvt =
1.897 m
Crvt
1.68 m
=
8
Cmvt =
1.3 m
Ctvt
=
0.84 m
Yvt
=
0.843 m
Performance Take off performance
The take off distance is given by Sto = (1.44×W2)/ (g×ρ×S×CLmax[T-(D+μ (W-L))] CLtakeoff = CLmax (Vstall/Vtakeoff )2 Assuming takeoff speed is 10 % greater than the stall speed CLtakeoff = CLmax (1/1.1)2 = 1.945 × 0.826 = 1.61 L= 0.5 × ρ×V2×S×CL = 0.5×1.2256×55.555 2×30×1.61 L= 91.35 KN We know that W= 110362.5 N and T/W=0.33 Hence T=36419.6 N Assuming L/D = 15 D=91350/15 = 6090 N Thrust required = 40 KN Therefore Sto = (1.44×W2)/ (g×ρ×S×CLmax [T-(D+μ (W-L))] Sto = 0.745 km Landing performance:
8
The landing distance is given by SLo = (1.69×W2)/ (g×ρ×S×CLmax (D+μ (W-L))) Here WLO = Wg-0.3 Wg = 77.25 KN Therefore SLo = (1.69×W2)/ (g×ρ×S×CLmax (D+μ (W-L))) SLo = 2.1 km Climbing flight:
R/c = V sin γ = V (T-D)/W = (PA-PR)/W R/C = 68.28 m/s 68.28= V sin γ γ = 17°53°
Horizontal turn:
Bank angle is Ø Turn angle is θ The sustained turn is considered for this: nm = (T/W) × (L/D) = 4.95 nmax is the maximum sustained load factor. But some factor of safety should be given hence n is assumed to be 3.5
8
L cos Ø = W Sec Ø = nmax Ø = sec-1(4.95) Ømax = 78°30° For normal horizontal turn n = 2.5 Hence Ø = 70°31° Tan θ = (n2 – 1)0.5 = 4.85 (n^2 – 1)0.5 = V2/gR2 = 4.85 There fore R = 1038.43 m. Radius has to be decreased for better performance. It is decreased by decreasing the velocity. ω = V/R = 0.21 m/s
Endurance:
The endurance for a jet engine is given by E= (L/D) × (1/C) ×ln (Wi/Wf) E = 9.59 hours.
Serial no 1 2 3 4
characteristics Take off distance Landing distance Rate of climb Climb angle
8
value 0.745 km 2.1 km 68.28 m/s 17°53°
5
Turn radius
1038.43 m
AIRCRAFT DATA SHEET 1
Cruise velocity
2
Design range
3
Cruise altitude
4
Aspect ratio
7.5
5
Sweep angle
31
6
Thrust /weight ratio
.33
7
Wing loading
315 Kg/M2
8
Max take off weight
11250 kg
9
Wing area
30
10
MTOW (N)
112435.6
11
Wpay (N)
32000
12
Wst (N)
35316
13
Wfe (N)
5513
14
Wpowerplant (N)
6396
15
Wfuel (N)
33211
16
Required Thrust (kN)
37
17
Thrust per engine (kN)
800 Km/hr 3825 Km 12750M
20
78
Cruise altitude (m)
11750
19
Sigma
0.26
8
20
Range (km)
3825
21
Cruise Velocity (kmph)
800
22
Time of cruise (hrs)
10
23
Power plant chosen
CF700-202
24
Dry weight (kg)
349
25
Net engine weight (N)
3423.7
26
Thrust per engine (kN)
20
27
Net Thrust (kN)
40
28
Cruise SFC (lb/hr/lbt)
0.65
29
Calculated W fuel (N)
3375
30
Aspect Ratio(AR)
7.5
31
Area(S)
30M2
32
Span(b)
15M
33
C r
3.053M
34
C t
0.947M
35
λ
0.31
36
Λc/4
29.5˚
AIRFOIL SELECTION
8
parameter
root
mean
Tip
Chord (m)
2.666
2.0
1.333
NACA
652415
64A410
65206
t/c
15 %
10 %
6%
Clmax
1.6
1.625
1.1
Cdomin
0.004
0.00425
0.0035
3-D VIEW DIAGRAM
1.74M 16M
ZZZZZZZZZZZZZZZZZZZZZZZZZ
ZZZ
8
30 M2
15 M
Conclusion The aircraft is designed and the parameters like cruise velocity, wing loading, span etc have been selected for our aircraft. The weight estimation had been done to estimate the weight of our aircraft. The wings, airfoil, landing gear have been selected for our aircraft. The performance calculations were also made to estimate the performance. The aircraft 8
parameters are in the optimum range and design characteristics have been found to be satisfactory
BIBILOGRAPHY 1.
“AIRCAFT DESIGN” by
PERFORMANCE
AND
John D.Anderson, Jr 2. “AIRCRAFT DESIGN” by Daniel P. Ramer 3. “ALL THE WORLD’S AIRCRAFT “ by janes 8
