Design project I HINDUSTAN COLLEGE OF ENGINEERING 10 SEATER BUSINESS JET
THRUSTMASTER-I
By, (reg no: 30507101064) (reg no: 30507101075) (reg no: 30507101306)
Acknowledgement
I would would like to extend extend my heart heartful ful thanks thanks to Dr. P.K. P.K. Dash, Head Head of Aeronautical Department for giving me his able support and encouragement. At this juncture I must emphasis the point that this desi design gn proj projec ectt woul would d not not have have been been poss possib ible le with withou outt the the high highly ly informative and valuable guidance by senior Prof. P.S. Venkatnarayanan, whose vast knowledge and experience has made us go about this project with great ease. We have great pleasure in expressing our sincere and whole hearted gratitude to them. It is worth mentioning about my teammates, friends and colleagues of the aerona aeronauti utical cal depart departmen mentt for extend extending ing their their kind kind help help whenev whenever er required. I thank one and all who have directly or indirectly helped me in making yhis design project a great success.
INTRODUCTION
Business jet, private jet or, colloquially, bizjet is a term describing a jet aircraft, aircraft, usually of smaller size, designed for transporting groups of business people or wealthy individuals. Business jets may be adapted for other roles, such as the evacuation of casualties or express parcel deliveries, and a few may be used by public bodies, governments or the armed forces. forces. The more formal terms of corporate of corporate jet , executive jet , VIP transport or transport or business business jet tend jet tend to be used by the firms that build, sell, buy and charter these aircraft.
The business jet industry groups the jets into five loosely-defined classes:
Heavy jets The most exclusive type of private jet is the heavy jet type, which is designed for the ultimate in large capacity luxury air travel. Aircraft of this class includes: •
Airbus Airbus A318 Elite Airbus A319CJ o
o
•
Boeing Boeing Business Jet
o
Large size jets •
Bombardier Aerospace Bombardier Global 5000 Bombardier Global Express o
o
Super mid-size jets The elite class of the business and private jet aircraft are the super mid-size jets that feature wide body cabin space, high altitude, speed, and ultra long range capabilities. These ultra luxurious private jets combine the long range transatlantic capability with the speed and comfort of a wide body, high altitude aircraft. Aircraft of this class include:
•
Bombardier Aerospace Bombardier Challenger 300 Challenger 605 o
o
Mid-size jets These aircraft are suitable for longer range travel such as transcontinental flights and for travel with larger passenger capacity requirements. Aircraft of this class includes: •
Bombardier Aerospace Learjet 60 XR Learjet 85 Gulfstream Gulfstream 150 Gulfstream 250 o
o
•
o
o
Light jets The light jets have been a staple of the business jet industry since the advent of the Lear Jet in the early 1960s. The light jets provide access to small airports and the speed to be an effective air travel tool. Aircraft of this class includes: •
Bombardier Aerospace Learjet 40 Learjet 40 XR Learjet 45 o
o
o
Very light jets Very light jets, jets , also known as Micro jets or VLJs, are designed to provide air travel, for example, to the more than 5,000 small community airports in the Unit United ed Stat States es.. VLJs VLJs have have a maxi maximu mum m take takeof offf weig weight ht of not not more more than than 10,000lb. Aircraft of this class include: •
Adam Aircraft Industries Adam A700 o
DESIGN PROCESS
A
irplane design is the intellectual engineering process of creating on
paper or in a computer aided screen , a flying machine to (1) meet certain spec specif ific icat atio ions ns and and requ requir irem emen ents ts esta establ blis ishe hed d by pote potent ntia iall user userss (or (or as perceived by the manufacturer) and/or (2) pioneer innovative, new ideas and technology. An example of the former is the design of the most commercial tran transp spor orts ts,, star starti ting ng at leas leastt with with the the Doug Dougla lass DC IN 1932 1932,, whic which h was was designed to meet or exceed various specifications by an airline company. An example of the latter is the design of the rocket powered Bell X-1, the airplane designed to exceed the speed of sound in the level of climbing flight. The design process is in fact an intellectual activity, but a rather a special one that is tempered by good intuition developed through experience, by attention paid to successful airplane design that have been used in the past, and by (gener (general al propri propriet etary ary)) design design proced procedure uress and databa databases ses (handb (handbook ooks, s, data data books etc.) that are a part of every airplane manufacturer.
Conceptual Conceptual Design Studies The first activity in the project design process is the conceptual design study. In this this phas phasee conv conven enti tion onal al and and nove novell conf config igur urat atio ions ns are are cons consid ider ered ed to determine layouts which are technically feasible and commercially viable, at the start of the phase all options are considered and during the conceptual design phase the quantity of data generated on each design will be relatively limited. The outcome of the study is the knowledge of the feasibility of various various concepts concepts and an estimate estimate of the likely likely dimension dimension and most favoured favoured configuration. Conc Concep eptu tual al desi design gn begi begins ns with with a set set of requ requir irem emen ents ts put put forw forwar ard d by a prospective customer or by any organization. The first preference will be given to the customer’s needs. Design details include aircraft range, payload, speed, versatility etc.
The actual design usually begins with a conceptual sketch. The conceptual sketch will include approximate wing and tail geometries, fuselage shape, locations of the engines, landing gear etc.
Preliminary Design Studies The The conc concep eptu tual al desi design gnss crea create ted d will will be comp compar ared ed and and thos thosee whic which h are are cons consid ider ered ed unfe unfeas asib ible le or too too comm commer erci cial al risk risky y will will be elim elimin inat ated ed.. The The remainder will be compared after careful consideration of the demands laid out by the customer. It is important not to carry many options forward to the next next stag stagee as this this will will diss dissip ipat atee the the avai availa labl blee effo effort rt and and slow slow down down the the detailed definition of the preferred design. However, care must be taken to avoid discarding design layouts too quickly as some may lead to evolutionary configurations which could give the aircraft a competitive advantage over aircraft from other companies.
Detailed Design Studies The The deta detail iled ed desi design gn phas phasee is star starte ted d towa toward rdss the the end end of the the prel prelim imin inar ary y analysis. In this part of the design process the layout is refined to a greater level of detail. With the external shape fixed the structural frame work will be defined. In this phase there will be an increasing reluctance to make radical geometrical changes to the overall layout of the aircraft. Throught this phase the aircraft weight and performance estimates will be continuously updated as more details of the aircraft layout becomes available.
SEVEN INTELLECTUAL POINTS FOR CONCEPTUAL DESIGN:
ABSTRACT
The following design requirements and research studies are set for the project: •
Design an aircraft that will transport 10 business class passengers and their baggage over a design range of 3500 km at a cruise speed of about 800 km/h.
•
To provide the passengers with high levels of safety and comfort.
•
To operate from regional airports.
•
•
To use advanced and state of the art technologies in order to reduce the operating costs.
To offer a unique and competitive service to existing scheduled operations.
•
To assess the development potential in the primary role of the aircraft.
•
To produce a commercial analysis of the aircraft project.
ALTERNATIVE ROLES
Our aircraft will have a fuselage size that is more spacious than normally associated with a 10-seat business jet. The long range requirement will demand a high fuel load and this will make the aircraft maximum design weight heavier than normal for 10 seat business class jet aircraft. Both this aspect suggests that the aircraft can be converted into a higher capacity short range airliner. A study will be required to investigate such varian variants. ts. This This type type invest investiga igatio tion n may result result in recomm recommend endati ations ons to change the baseline aircraft geometry to make such developments easier to achieve. For example, using the extra fuselage space up to six abreast seating in the higher capacity aircraft can be made. Without such a change, six abreast seating may not be feasible.
COMPARITIVE DATA SHEET
To have an idea about the parameters for our design process we compare the important parameters of already existing aircraft similarly related to our interest. The important parameters thus obtained are plotted with respe espect ct to the the crui cruisi sing ng spee speed d and and the appr pproxim ximate ate value alue of the parameters for the new aircraft is found out.
The parameters involved in the CDS are listed below: Aspect
ratio
Maximum takeoff weight Empty weight Power Cruise
speed
Wing loading Range
Length
Wing span
Wing area
S No
NAME OF AIRCRAFT
CAPACIT Y
ALTITUDE
RANGE
SPEED
ASPECT RATIO
(passenge rs)
(m)
(km)
(km/h)
1 4 +2
12800
3 350
8 62
6 .5
1
Falcon 20F
2
Piaggio P 180 Avanti
9 +2
12500
2592
7 32
12.97
3
Cessna Citation I
7 +2
12450
2559
9 16
7 .3
1 3 +2
12500
4 893
8 30
7.003
6 +2
5914.8
1 9 15
3 46
5.31
7 +2
13720
2037
8 86
9.765
1 0 +2
13458
8 790
8 90
9 .9
4
5
Hawker 850 XP Aero commander 500 S
6
Learjet 24D
7
Raython Hawker 800 XP
8
A38
8 +2
4800
4 0 59
7 90
7 .4
9
Bombadier Learjet 40
7 +2
15540
3156
8 65
9 .9
10
Bombadier Learjet 60R
9 +2
15545
4330
8 59
12.99
11
Challenger 800
1 5 +2
12497
5 129
8 50
14.82
12
Bombadier Global XRS
1 2 +2
15500
11390
8 60
9 .6
13
Global 5000
1 3 +2
15554
9 360
9 50
15.3
14
Dassult Falcon
1 4 +3
15542
11019
9 53
9.709
15
Learjet 24E
7 +2
13724
2052
8 86
8.58
S N o
NAME OF AIRCRAFT
WING LOADING
WING SPAN
LENGTH
MAX TOW
R/C
WING AREA
(km/sq.m)
(m)
(m)
(kg)
(m/min)
(sq.m)
1
Falcon 20F
3 54
16.3
17.15
13000
1524
41
2
Piaggio P 180 Avanti
3 27
14.03
14.41
5239
900
16
3
Cessna Citation I
2 08
14.35
13.26
5375
828
25.9
4
Hawker 850 XP
4 41
16.5
1 5 .6
12701
870
34.75
5
Aero commander 500 S
4 92
14.95
11.22
3060
409
23.7
6
Learjet 24D
2 99
13.12
1 4 .5
6136
2073
21.53
7
Raython Hawker 800 XP
3 90
14.56
16.39
9545
859
32
8
A38
4 71
16.6
1 2 .8
3150
6 96
67
9
Bombadier Learjet 40
3 90
14.56
16.93
9545
859
28.95
1 0
Bombadier Learjet 60R
4 33
13.34
17.88
16600
6 60
24.6
1 1
Challenger 800
4 51
21.21
26.77
24041
7 56
48.35
1 2
Bombadier Global XRS
4 68
28.65
3 0 .3
44500
793
94.9
1 3
Global 5000
4 51
28.65
2 9 .5
41957
810.6
53.29
1 4
Dassult Falcon
4 35
26.21
23.19
31750
774.6
70.7
1 5
Learjet 24E
3 42
13.35
14.52
6136
850
24.57
DESIGN DATA
PARAMETER
VALUE
MAXIMUM SPEED
720km/h 10.25 20m 22.88m 16,642kg 6000km 20,000kg 10200m 10 98KN 750km/h
WING LOADING
350kg/m2
CRUISE SPEED ASPECTRATIO LENGTH WING SPAN WEIGHT RANGE MAX. TAKEOFF WEIGHT SERVICE CEILING CAPACITY THRUST
WEIGHT ESTIMATION
OVERALL WEIGHT ESTIMATION
The The sec second ond pivot ivot poin pointt in our con concept ceptua uall desi desig gn anal analy ysis sis is the preliminary weight estimation of the aircraft. There are various ways to subdivide and categorize the weight components of the airplane. 1.
Crew weigh Wcrew : The crew comprises of the people necessary to operate the aircraft.
2.
Payl Payloa oad d weig weight ht W payload : The The payl payloa oad d is what what the the airp airpla lane ne is intended to transport i.e. passengers, baggage etc.
3.
Fuel weight Wfuel : it is the weight of the fuel in the fuel tank. Since the the fuel fuel is cons consum umed ed duri during ng the the cour course se of the the flig flight ht,, Wfuel is variable, decreasing with time during the flight.
4.
Empty weight Wempty : This is the weight of everything else the structure, the engines, electronic equipments, landing gear, fixed weights etc.
The design take-off weight is the weight of the aircraft at the instant it begins its mission. It includes the weight of all the fuel on board at the beginning of the flight.
W=WSTRUCT+WFIXEDEQUI+WCREW+WPASSENGER +W +WPOWERPLANT+WFUEL
W fuel is approx. =0.3Wtakeoff Wstructure is approx. =0.32 Wtakeoff W power plant is approx. =0.055 Wtakeoff Wfixed equipment is approx. =0.05 Wtakeoff Wcrew+W passengers is approx. =0.275 Wtakeoff
WSTUCT
=0.32 Wtakeoff
WSTUCT
=6432kg
WFIXED EQUIP
=0.05 Wtakeoff
WFIXED EQUIP
=1005kg
WPOWERPLANT
=0.055 Wtakeoff
WPOWERPLANT
=1105.5kg
WFUEL
=0.3 Wtakeoff
WFUEL
=6030kg
WCREW+WPASSENGER
=0.275 Wtakeoff
WCREW+WPASSENGER
=5527.5kg
W=0.725W+(no. Of passengers)*180+(crew)*100 passengers)*180+(crew)*100 =0.725(20,000)+((10*1800)/9.8) =0.725(20,000)+((10*1800)/9.8)+((3*1000)/9.8) +((3*1000)/9.8) =18,142Kg Therefore the gross weight of the aircraft is 18,142 Kg
AEROFOIL SELECTION AIRFOIL TERMINOLOGY The various terms related to airfoils are defined below :
The mean camber line is a line drawn midway between the upper and lower surfaces
The cord line is a straight line connecting the leading and the trailing edges of the aerofoil, at the ends of the mean caber line.
The chord is the length of the chord line and is the characteristic dimension of the airfoil.
The maximum thickness and the location of it are expressed as a percentage of the chord.
For the symmetrical aero foil both mean camber line and chord line pass from the centre of gravity of the aero foil and they touch at leading and trailing edge of aerofoil.
The The aero aerody dyna nami micc cent centre re is the the chor chord d wise wise leng length th abou aboutt whic which h the the pitching pitching moment is independe independent nt of the lift coefficien coefficientt and the angle angle of attack.
The The cent centre re of pres pressu sure re is the the chor chord d wise wise loca locati tion on abou aboutt whic which h pitc pitchi hing ng moment is zero. NACA airfoils are airfoil shapes for aircraft wings developed by national advisory advisory committee committee for aircraft aircraft wings develo developed ped by the Nation National al Adviso Advisory ry Committee for Aeronautics “NASA”. The shape of the NASA airfoils is desc descri ribe bed d usin using g a seri series es of digi digits ts foll follow owin ing g the the word ord “NAC “NACA” A”.. The The parameters in the numerical code can be entered into equations to precisely generate the cross section of the airfoil and calculate its properties.
CRUISE VELOCITY= 720 km/h=200m/s ASPECT RATIO
= 10.25
CRUISE ALTITUDE=10200 m From the graph we have W/S= 350 kg/m 2 CL=2×(W/S)/ρv2stall VSTALL=0.25Vcruise=180 km/hr CL= 0.17 Thus CL was found to be 0.17. With this as the primary criterion for airfoil selection, we opt for the airfoil NACA 63-210. THICKNESS TO CHORD RATIO:
For For lowlow-sp spee eed d subs subson onic ic airc aircra raft ft rela relati tive vely ly high high t/c t/c valu values es (up (up to 0.2) 0.2) acceptable at wing root – gives good structural depth with small profile drag penalty.From the last two digit NACA aerofoil we have a thickness to cord ratio of about 10% of the chord. This aerofoil was selected selected due to its apt t/c ratio of around 10 % of the chord and has a design lift coefficient as required. The details of the aerofoil are:
The 2nd digit indicates the distance of min. pressure area i.e. 30% of the chord. The 3rd digit indicates the design lift coefficient in tenths.C L=0.2. The final 2 digit indicates the maximum thickness in tenths of % of the chord. In this case it is 10%.
NACA 63-2100
DRAG POLAR
ON THE SELECTED AIRFOIL
CHARACTERISTICS OF AIRFOIL
WHEN KEPT IN A WIND TUNNEL
ENGINE SELECTION
The gross weight estimate estimate of the aircraft aircraft has been done. done. It is known that the weight of the power plant used is roughly 5.5% of the overall weight.The gross weight of the aircraft is given by Wo=18,142 kg From this it is clear that the power plant weight is approximately, Wpower plant =0.055*18142=998kg =0.055*18142=998kg In most of business jets, it is common practice to have twin engine and our design has the same configuration. Hence the weight of the engine is 500kg. From the design sheet it was known that that the engines should produce approximately approximately 98 KN of thrust ,i.e.= 49 KN ≈ 50 KN each. The table below lists few engines that meet our demands. COMPANY
ENGINE
THRUST(KN)
WEIGHT
SFC
RB 183-2 Mk
45.8
716
0.75
CF 34 3A
46.8
520
0.42
CF 34-3B
44.2
467
0.45
Rolls Royce General Electric General Electric Pratt & Whitney
PW810
48
512
0.62
Rolls Royce
AE3007A1P
54
478
0.59
From the above chart we select for Rolls Royce AE3007 A1P.
Rolls Royce AE 3007
Fuel weight validation
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 1.2 Cruise velocity The factor of 1.2 is provided for reserve fuel.
Thrust at altitude is calculated using the relation
T σ
1 .2
=
T 0 * σ
Tσ is the thrust at the altitude and T o is the thrust at sea level. σ
ρ a lt
=
ρ 0
ρalt=density of air at 10000m and ρ o is the density of air at sea level. Atmospheric Chart Pressur Tem Altitude e p. (feet) (in. Hg) (F°)
Densit y (%)
sea level
29.92
59.0
100
2,000
27.82
51.9
94.3
4,000
25.84
44.7
88.8
6,000
23.98
37.6
83.6
8,000
22.22
30.5
78.6
10,000
20.57 20.57
23.3
73.8
12,000
19.02 19.02
16.2
69.3
14,000
17.57
9.1
65.0
16,000
16.21
1.9
60.9
Ther Theref efor oree from from the the abov abovee char chart, t, at 10,0 10,000 00 m the the dens densit ity y decr decrea ease sess by 26.2%.Therefore the density becomes 0.85 kg/m 3 . σ=0.85/1.225=0.75 Tσ=98*0.751.2=34.69 Wfuel = (no. of engines) x (thrust (thrust at altitude) altitude) x Range x SFC SFC x 1.2 Cruise velocity Wfuel = 2*34.69 34. 69*1 *100 000* 0*6000 6000*0. *0.8* 8*1.2 1.2 720 Wfuel=5600kg
WING DESIGN
The design design of wing wing involve involvess buildi building ng a wing wing with apt shape shape and the struct structure ure to impro improve ve aerody aerodynam namic ic effici efficienc ency. y. The wing wing is basica basicall lly y extrusion of the 2D airfoil with a sweep. For a subsonic aircraft it is best to have the root airfoil 20-60% thicker than the tip airfoil as it reduces weight and gives more volume for fuel and landing gear. Wing area(S)=(weight/wing loading) =20100/350 S =57.42m2 A.R=b2/s =10.25(from design data sheet) From this we find the wing span(b)=24.26m Croot =(2S/b(1+λ)) Taper ratio(λ)=Ctip/Croot=0.3 for jet aircrafts
Croot = 3.7m Ctip = 1.11m`
DETERMAINING THE MAC: Cmean= [0.66*Croot*(1+λ+λ 2/(1+λ))] Cmean= 2.61m
Distance of mean chord from aircraft centre line = [b*(1+2λ)/6(1+λ)] [b*(1+2λ)/6(1+λ)] = 4.97m
Wing sweep has stability. A swept wing has natural dihedral effect. Sweep back angle at leading edge[ ʌLE] =300
LOCATION OF THE WING: As in the case of business jets the wing is of low wing configuration. The primary advantage is that it helps in stowage of landing gear.
HIGH LIFT DEVICES: Flaps are high devices which, in effect ,increase the camber of the wing and, in some cases, as with the fowler flap, also increase the effective wing area. Their use permits better takeoff performance and permits steeper approach angles and lower approach and landing speeds. When When defl deflec ecte ted, d, flap flapss incr increa ease se the the uppe upperr cham chambe berr of the the wing wing,, the the negative negative pressure pressure on the wing. wing. At the same time, they allow allow a build up of pressure below the wing.
During takeoff, flap settings of 100 to 200 are used to better take off per perfo form rman ance ce and and a bett better er angl anglee of clim climb, b, espe especi cial ally ly valu valuab able le when when
climbing out over obstacles. Flaps do indeed increase drag. The greater the flap deflection, the greater the drag. At a point of about half of their full travel, the increase drag surpasses the increase lift and flaps become airbrakes. Most flaps can be extended to 400 from the chord of the wing. At settings between 200 to 400,the essential function of the flaps is to incr increa ease se the the land landin ing g capa capabi bili liti ties es,, by stee steepe peni ning ng the the glid glidee with withou outt increasing the glide speed. In an approach over obstacles, the use of flaps permits the pilot to touch down much nearer the threshold of the runway. Flaps also permit a slow landing speed and act as airbrakes when the airplane is rolling to a stop after landing, thus reducing need for excessive breaking action. As a result, there is less wear on the under carr carria iage ge,, whee wheels ls and and tyre tyres. s. Lowe Lowerr land landin ing g spee speeds ds also also redu reduce ce the the possibility of ground looping during the landing roll
Slotted flaps, on the other hand, including such types as fowler and a zap produces lift in excess of drag and their partial use is therefore recommended for takeoff.
From From the the stan stand d poin pointt of aero aerody dyna nami micc effi effici cien ency cy,, the the fowl fowler er flap flap is gene genera rall lly y cons consid ider ered ed to offe offerr the the most most adva advant ntag ages es and and the the fewe fewest st disadvantages, especially especially on larger larger airplanes, while while double slotted flaps flaps have won wide approval for smaller types.
Chan Change gess in flap flap sett settin ing g affe affect ct the the trim trim of an airp airpla lane ne.. As flap flapss are are lowered the centre of pressure moves rearward creating a nose down pitching moment. However, in some airplanes ,the change in air flow over the tail plane as flaps are lowered, is such that the total moment created is nose up and it becomes necessary to trim the airplane nose down.
The airplane is apt loose considerable height when the flaps are raised. At the low altitudes, therefore the flaps should be raised cautiously.
Fig. Location of flaps and ailerons in the wing
FUSELAGE CABIN LAYOUT
The cabin cabin layout of the business business jet aircraft aircraft is designed for comfort comfort rather rather than seating capacity. The cabin may also have a separate section for the meeting and the refreshment centre. The design parameters of the cabin layout are: Seat pitch
=1.02m
Seat width
=0.74m
Aisle width
=0.6m
Cabin diameter
=2.16m
By standards described by the Raymer, the structural thickness is given by: t=0.002dfus+0.0254m=0.0686m The external diameter is thus =2.1+(0.0686*2)=2.2972m
The total length of the fuselage = 23 m.
As per raymer suggestion the two crew aircraft should have a minimum length of 100 inches nose. So we choose the length of the nose to be 20% of the total length. i.e. 0.2(23)= 4.5m. The length of the tail is taken as about 25% of the length which is equal to 5.6m. The main reason for choosing such a high value for the tail is explained below. Having a highly inclined tail may increase the angle of attack of the wing during the takeoff in or near to the ground. The high sweep angle of the tail may raise the rear fuselage level giving enough angle to the wing.
DETERMINATION OF CENTRE OF GRAVITY
Considering forces to act at the centre of various sections and taking moments about the nose C.G can be determined. Hence,
C.G. = (160*2.25)+(1240*10.95)+(85.75*1 (160*2.25)+(1240*10.95)+(85.75*10.9)+(1105*18) 0.9)+(1105*18) 160+1240+85.75+1105 C.G.=13.41m
LANDING GEAR DESIGN The undercarriage or the landing gear is the structure(usually wheels)that supports an aircrafton the ground and allows an aircraft on the ground and allow it to taxi.
Types of landing gear: Wheeled under carriage comes in two types: Conventio Conventional nal or tail drageer drageer under carriag carriagee ,where there there are two main wheels towards the front of the aircraft and a single ,much smaller ,wheel or skid at the rear. Tricycle under carriage ,where there are two main wheels under the wing wingss and and a third ird smal smalle lerr whee wheell in the the nose nose.. The The tail ail dragg agger arrangement was common during the early propeller era,as it allows more clearence for the propeller clearence . Most of the modern aircraft have tricycle arrangement.tail draggers are cons consid ider ered ed hard harder er to land land and and take take off( off(be beca caus usee the the arra arrang ngem emen entt is unst unstab able le a smal smalll devi deviat atio ion n from from the the stra straig ight ht line line trav travel el is natu natura rall lly y amplified by the greter drag on the main wheel which has moved further away away from from the the plan plane’ e’ss cent centre re of grav gravit ity y due due to the the devi deviat atio ion) n),, and and usually require special pilot training.
Some Someti time mess a smal smalll tail tail wheel wheel or skid skid is adde added d to the aircr aircraf aftt with with tric tricyc ycle le unde underc rcrr rria iage ge,i ,in n case case of tail tail stri strike kess duri during ng the the take take off. off. The The conocorde,for instance,had arectractable tail bumper wheel. The Boeing 727 also had a rectractable tail bumper. Some aircraft with rectractable conven conventio tional nal landin landing g gear gear having having a fixed fixed tail tail wheel, wheel,whi which ch genera generate te minimum drag( since most of the airflow past the tail wheel has been bla blank nket eted ed by the the fuse fusela lage ge)) and and even even impr improv ovee yaw yaw stab stabil ilit ity y in some some cases.
Rectractable landing gear: To decrease drag during the flight some under carriages retract into the wings and /or concealed behind doors;this is called rectractable landing gears. A design for retractable landing gear was first seen in 1876 in plans for an amphibious monoplane designed by Frenchmen Alphonse and Paul Gauchot. Aircraft Aircraft with at least partial partially ly retractable retractable landing landing gear did not appear until until
1917, and it was not until the 1920s and early 1930s that such aircraft became common. By then then,, airc aircra raft ft perf perfor orma manc ncee was impr improv oved ed to the the poin pointt wher wheree the the aero aerody dyna nami micc adva advant ntag agee of a retr retrac acta tabl blee carr carria iage ge just justif ifie ied d the the adde added d com complex lexity ity and and weigh ight. An aero erodyn dynami amic appr approa oach ch of reduc educin ing g the aero aerody dyna nam mic penal enaltty impo impose sed d by fixed ixed unde underc rcar arri riaage is to attach tach aerodynamic fairings (often called spats ) on the undercarriage, with only the bottom of the wheels exposed. Pilots Pilots conforming conforming that that their landing landing gear is down refer refer to “three green” green” or “three in the green”, a reference to the electrical indicator lights from the nose wheel and two main landing gears. Amber Amber light light indicates indicates that the gears are in the up-locked up-locked position; position; red lights indicate indicate that the landing landing gear is in transit (neither (neither down down or locked locked nor fully retractable).
Landing gear placement:
The sizes of the tyres depend on the load distribution between the main wheels and the nose wheel. The load carried by each tyre is the equal and opposite reaction force exerted by the ground on them. F N and FM are the force acting on the nose and the main landing gear respectively. The takeof takeofff weight weight acts acts throug through h the centr centree of gravit gravity. y. The distan distance ce between the line of action of C.G. and the F N is given by X1.
The distance between the line of action of C.G. and the FM is given by X2.The distance between F N and FM is 9.5m.
Force diagram for obtaining load distribution among tyres.
Taking the moments at the respective points, we get FM= WoX1/X3 F N= WoX2/X3 We should note that FM is the sum of the reaction forces at the two main landing gears. The takeoff weight is 20,000kg.
FM= 20000*(7.9/9.5)=16631.5kg 20000*(7.9/9.5)=16631.5kg F N= 20000*(1.6/9.5)=3368.42kg Hence the load on the nose landing gear is 3368.42 kg. And the load on the main landing gear is 8315.75 kg each. RAYMER’s relations derived empirically for the diameter and width of tire are based on the load on the wheel. Wheel diameter or width of the wheel= AWB
A
B
Wheel diameter(in)
1.51
0.349
Wheel width(in)
0.715
0.312
Main wheels: Wheel dia. =A(FM/2)B =1.51(8315.75/2)0.349 =27.66in=0.702m
Wheel width= A(FM/2)B =0.715(8315.75/2)0.312 =9.62in=0.22m
Nose Wheels: Wheel dia. =A(F N/2)B =1.51(3368.42/2)0.349 =20.18in=0.51m
Wheel width= A(FM/2)B =0.715(3368.42/2)0.312 =7.26in=0.18m
BRAKING SYSTEM AND TYRES: The bias ply tyre,shown in the first picture consists of casing plies runnin running g diagon diagonall ally y at approx approxim imate ate right right angles angles to one anothe another. r. The number number of plies plies and the angles angles at which which they are laid laid dictate dictate strength strength and load capacity . the latest high performance bias ply aircraft tyres inter thread reinforcing fabric(ITF). This provide additional high speed stability reduces thread distortion
Under load,protects the casing plies from damage and can act as wear indicators or rethredable tyres. Bias tyres are currently the most popular tyres fitted to the world’s fleet.
The braking system used has an anti locking braking system,which stop the wheel heelss from lock ocking ing when when fully lly appl appliied ,all ,allow owiing grea greatter deceleration and control during breaking,particularly in wet conditions.the brakes,developed by Dunlop,were carbon based and could bring to a halt within a mile, and with proper cooling time to dissipate heat later.
PERFORMANCE ANALYSIS
TAKE OFF PERFORMANCE: Total take off distance=Sg+Sa Ground roll distance= 1.44*w^2{g*ρ*s*CLMAX(T-(D+µ(W-L)))0.7VLO}
Lift of velocity(VLO) =1.2*VSTALL =220 km/h 0.7 VLO=50.17m/s The factor due to ground effect is given by, Φ=16[h/b2]2/1+[16h/b] 2 h=1.92m b=1.85m Φ=0.04 For the dry surface, µ=0.04 D=3069.7N Sg=1.38km is the min. distance distance covered by the aircraft in the the ground before takeoff.
CALCULATION OF DISTANCE DISTANCE WHILE AIBORNE AIBORNE TO CLEAR AN OBSTACLE
Sa=R*sinΦOB R=V2stall*6.96/g = 25,638.3m Obstacle height is given by, hOB=10.668m(for jet) cosΦOB=(1-( hOB/R))
Flight path angle, ΦOB=cos-1(1--( hOB/R)) =1.74470
g-acceleration due to gravity
Airborne distance, Sa=R*sinΦOB=25,638.3*sin1.74470 =780.58m Total take off ditance=Sa+Sg =1.38+0.780 =2.2 km Therefore the aircraft requires a runway of 2.2 km to takeoff. CLIMB: R/C=Vstall(T-D)/W =20.833m/s R/C= Vstall sinΦ Φ=
sin-1(R/C/ Vstall)
Φ=
6.640
3D VIEW DIAGRAMS OF THE AIRCRAFT TOP VIEW
SIDE VIEW
FRONT VIEW
BIBLIOGRAPHY
1.AIRCRAFT DESIGN: A CONCEPTUAL APPROACH, DANIEL P. RAYMER, AIAA EDUCATION SERIES. 2.INTRODUCTION TO FLIGHT.J.D. ANDERSON 3.DESIGN OF AEROPLANE BY DARROL STINTON 4.www.aerospaceweb.org 5.www.wikipedia.org 6.www.airliners.net