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MINISTRY OF EDUCATION AND SCIENCE OF THE RUSSIAN FEDERATION SAMARA STATE AEROSPACE UNIVERSITY (NATIONAL RESEARCH UNIVERSITY)

CONCEPTUAL AIRCRAFT DESIGN
Term Project Example

SAMARA 2011

1

Compiler:

Reznichenko Gennadiy A.

Translated by: Kancher Galina S. Conceptual Aircraft Design = [Electronic resource]: Term Project Example / The Ministry of Education and Science of the Russian Federation, Samara State Aerospace University; Compiler G.A. Reznichenko. Electronic text and graphic data (345 kb). - Samara, 2011. - 1 CD-ROM. Guide is an example of a term project on discipline "Conceptual aircraft design" for the specialties 160201 "Airplane and Helicopter Construction", 160901 "Technical Maintenance of Aircrafts and Engines" and for the Masters programme лDesigning, construction and CALS-technologies in Aeronautical Engineering╗ for education direction 160100.68 лAeronautical Engineering╗. Prepared by the Department of Aeronautical Engineering SSAU.

й Samara State Aerospace University, 2011

2

CONTENTS 1 STATISTICS ANALYSIS ..................................................................................... 9 1.1 Scientific and technological forecasting ........................................................ 11 1.2 Analysis of the project situation .................................................................... 13 2 DEVELOPMENT OF INITIAL SPECIFICATION ............................................ 15 2.2 General technical requirements ..................................................................... 15 2.3 Aircraft performance requirements................................................................ 16 2.4 Production and technological requirements .................................................. 16 2.5 Performance Requirements ............................................................................ 16 2.6 The technical and economic requirements .................................................... 16 3 BASELINE CONFIGURATION DEVELOPMENT .......................................... 18 3.1 Selection of the wing parameters ................................................................... 18 3.2 Selection of the fuselage parameters ............................................................. 18 3.3 The choice of the tail parameters ................................................................... 19 3.4 Selection of the control surfaces parameters ................................................. 19 3.5. Selection of the undercarriage parameters.................................................... 19 3.6 The relative position of the aircraft units...................................................... 20 3.7 Selection of the propulsion system ................................................................ 20 3.8 Selecting the number of engines and their placement on the aircraft ........... 20 4 DETERMINATION OF INITIAL AIRCRAFT PARAMETERS ...................... 21 4.1 Determination of the specific load on the wing ............................................. 21 4.2 Aerodynamic parameters ............................................................................... 21 5 IDENTIFICATION OF AIRCRAFT THRUST NEEDS .................................... 22 6 DETERMINATION OF AIRCRAFT TAKEOFF WEIGHT .............................. 23 6.1 Definition of the initial take-off weight ......................................................... 23 6.2 Definition of the payload mass ...................................................................... 23 6.3 Definition of the equipment mass and service load ....................................... 23 6.4 Determination of the relative weight of construction .................................... 23 6.5 Determination of relative weight of the fuel system ..................................... 24 6.6 Determination of the relative masses of powerplant ..................................... 25 6.7 Determination of relative weight of the equipment and controls .................. 25 6.8 Definition of take-off weight of the first approximation ............................... 25

3

7 DETERMINATION OF THE ABSOLUTE DIMENSIONS OF THE AIRCRAFT ............................................................................................................. 26 7.1 Selection of engines ....................................................................................... 26 7.2 Determination of the fuel mass and volume .................................................. 26 7.3 Specifying the wing parameters..................................................................... 26 7.4 Definition of the plumage parameters ........................................................... 27 7.5 Determination of the fuselage dimensions .................................................... 27 7.6 Determination of the undercarriage parameters ............................................ 28 8 AIRCRAFT WEIGHT ESTIMATION ................................................................ 29 8.1 Wight summary.............................................................................................. 29 9. AICRAFT LAYOUT .......................................................................................... 30 9.1 Aircraft volume-weight assembly.................................................................. 30 9.2 The structural layout of the aircraft ............................................................... 30 10 AIRCRAFT BALANCE .................................................................................... 32 10.1 Choice of balance range............................................................................... 32 10.2 Specified loading conditions........................................................................ 33 11 DEVELOPMENT OF A GENERAL DRAWING VIEW AND AIRCRAFT TECHNICAL DESCRIPTION ............................................................................... 35 CONCLUSION ....................................................................................................... 36 BIBLIOGRAPHY ................................................................................................... 37

4

SYMBOLS , н volume coefficients of horizontal and vertical tail. н speed of sound; н wing angle of attack; B, н landing gear track, relative landing gear track; b н wing or tail chord; b н mean aerodynamic chord of wing or tail; b н root chord of wing or tail; b н tip chords of wing or tail; н thickness-to-chord ratio of wing or tail; н specific fuel consumption of turbojet engine; н specific fuel consumption of turboprop/turbofan engine; , н drag and lift coefficients in the velocity-related coordinate system; Cxa0 н zero-lift drag coefficient (y=0); D н induced drag ratio; D н fuselage diameter;

н deflection angle of control surfaces and wing high-lift devices;
f н safety factor, the friction coefficient; g н acceleration due to gravity;

н landing gear offset angle; н engine specific weight;
н flight altitude; н sweep angle of wing or tail; н lift-to-drag ratio; н coefficient; L н flight range; l н wing span, tail span; l

н the takeoff run;

н aspect ratios of wing or tail;
5

н Mach number; m н weight of an aircraft or its part, engine bypass ratio; N н engine power; N н aircraft power-to-weight ratio; n, n н ultimate and limit load factors; n

н number of passengers;

P н engine thrust; P0 н aircraft start thrust-to-weight ratio; 0 н wing loading; q н dynamic pressure;

н air density;
н air relative density; S н wing area, tail area; S н the relative tail area; н wing taper ratio, tail taper ratio; V н flight velocity; Vy н rate of climb; X н center of gravity coordinate; XF н aerodynamic center coordinate;

н aircraft overturning angle; н coefficient to account for thrust change with flight altitude; н engine throttling coefficient; н aircraft ground angle; н coefficient to account for thrust change with flight velocity.

6

ABBREVIATIONS daN=10N н decanewton; MAC н mean aerodynamic chord; TJEA н turbojet engine with the afterburner duct; BTJE н the bypass TJE; RE н the reciprocating engine; TPE н the turbo-propeller engine; TPFE н the turbo-propeller and fan engine;; SUBSCRIPTS н wave; н takeoff; / н vertical/horizontal empennage; н engine; н landing approach; , н frame, airframe; н commercial; н ruise, ruising; н critical, limit; 0 н initial, start value; н equipment and control system; н airplane takeoff; н climb; н air navigation margin; н passengers; н landing; н payload; н total load; , () н altitude limit; н empty; 7

н calculated; н takeoff run; н flight; / н elevator/rudder; н equipment; н power plant; н fuselage; н payload; н landing gear; н crew.

8

1 STATISTICS ANALYSIS For the development of initial specification, you must first examine the statistical material, for which we choose the five planes of different schemes: the normal scheme. Boeing 747-400 Boeing Model 747-400 - long-haul passenger aircraft developed by the American company Boeing. Airbus A330-300 (Airbus A330-300) A330-300 - A340-300 prototype, but has only 2 engines. Range up to 8980 km. Airbus A380 Airbus A380 - wide-body passenger jet, created concern лAirbus SAS╗. Table 1- Basic Data of the Prototype Aircraft
1 Aircrafts Airplane model, manufacturer, year and country Crew, people Characteristics of the propulsion system Engine type, number (n), thrust 0 (daN), power N0 (kW) 2 Airbus A340-300, Boeing 747Airbus SAS, 400 Freighter, European Union, Boeing, USA, 1992 1989 2 2 Turbofan PW4062, 4 x 28600 Turbofan CFMI CFM56-5C4, 4 x 15400 1 3 4 5 Airbus A330Il-96-400M, 300, Airbus SAS, Ilyushin EDB, Airbus A380, European Union, Russia, 2001 European 1992 Union, 2005 2 Turbofan P&W PW4168A, 2 x 30800 2-3 RDD PS-90A-1, 4 x 17400 2 Rolls-Royse Trent 900 or Engine Allianse GP 7000 4 x 31100 0 ,5 5 7 8 0 ,2 3 1

2 3

4 5 6

7 8 9 10 11

Fuel consumption rate C0 (kg/daNh), 0 (kg/kWh) By-passratiom Specific engine weight e, ( = meg, daN/kW) Weight characteristics Take-off mass m0, kg Payload (combat load) mass m, kg Empty airplane mass m, kg Fuel mass m, kg Specific wing load 0 , daN/m2

0 ,5 9 7 5 ,1 0 ,1 5 0

0 ,5 9 6 6 0,1140

0 ,5 9 1 6 0 ,1 3 4

0 ,5 8 0 4 ,8 0 ,1 8 3

396900 70620 181120 145160 0 ,8 2 2 0 ,1 9 4

275000 50900 130900 94000 0 ,8 1 5 0 ,1 8 8

230000 51700 122200 5800 0 ,7 7 5 0 ,2 2 5

270000 58000 123000 123622 0 ,7 8 5 0 ,2 1 5

560000 66400 276800 251100 0 ,8 8 1 0 ,1 0 2

9

12

Weight efficiency - = 0 or Thrust-to-weight ratio (power-to-weight 10 ratio) 0 = 0; (0 =
10 0 0 0 0 = 0

7 3 3 ,5

7 6 0 ,5

6 3 6 ,1

7 7 1 ,4

650.13

13

0 ,0 7 9

0 ,0 5 6

0 ,1 3 4

0 ,0 6 4

0 ,0 1

) (kW/daN)

14 15 16 17 18 19 20

Geometric characteristics Wing area S, 2 Wing span ,

5 4 1 ,1 64,4/7,66 9,54/3,66 3 7 ,5 0,09; 0,09 6,57/10,56 2/3,4

3 6 1 ,6 60,3/10,06 7,2/5 30 0,25; 0,125 5,64/11,28 1,7/3

3 6 1 ,6 60,3/10,06 7,2/5 30 0,25; 0,125 5,64/11,28 1,7/3

350 60,1/10,32 8,6/5 30 0,18; 0,1 6 ,0 8 2/3,2

845 79.8/7.54 11.9/4.37 3 3 .5 7 .1 4 2,1/3,6

21 22 23 24 25 26

Wing aspect ratio / wing taper ratio Wing sweep angle0 Relative thickness0 Fuselage diameter D, / fuselage aspect ratio Fuselage nose part aspect ratio / fuselage aft portion aspect ratio / Relative distance from fuselage nose to wing central chord Horizontal tail (HT) area S ,2/S HT aspect ratio / HT taper ratio / HT sweep angle HT arm L, m / Coefficient of HT static moment = Vertical tail (VT) S, m2 /S VT aspect ratio / taper ratio / VT swept angle

0 ,2 5 8 1 2 7 ,9 3,75/3,28 35 38,15/3,99 0 ,5 4 3 8 0 ,6 4 4,57/2,25 30 29.79/4,13 0 ,5 9 5

0 ,3 5 8 0 .6 4 4,57/2,25 30 29.79/4,13 0 ,5 9 5

0 .2 9 6 9 3 .6 5 4.5/4 30 29.594/4.33 0 ,9 2 0

0 .3 3 2 231 4.3/2.5 35 44.03/3.69 0 .7 2 5

27 28 29 30

area VT

115 1,19/3,37 45 36,05/0,56

5 0 ,4 1,4/3,2 45 30,51/0,505

5 0 ,4 1,4/3,2 45 30,51/0,505

51 1.46/2.7 45 27.773/0.46

125.03 1.68/3.89 45 36.260/0.45

VTarmL m /

=

31

Coefficient of VT static moment Landing gear tread

0 ,2 6 1

0 ,1 6 3

0 ,1 6 3

0 .5 9

0 .4 5

32

=

33

34

Main landing gear offset Performance characteristics Maximum velocity over altitude Vmax/H, km/(h*m)

=

0 .3 6 2

0 ,4 0 3

0 ,3 8 2

0 .4 0 8

0 .4 1 6

0 .1 7 1

0 ,1 7 7

0 ,1 7 7

0 .1 7 3

0 .2 9

1150/10670

1100/12500

1050/11800

900/12000

950/15200

10

35 36 37 38 39 40 41 42 43 44 45 46 47 48

Cruise velocity over altitude V/H, km/(h*m) Landing velocity V(V) , kmph Flight range with full payload L, km Flight range with reduced payload Lmax, km Take-off run (or runway length L), m Climb rate Vy0, m/s Maximum altitude , m Others Number of passengersn Cargo compartment dimensions BxHxL, m mm Airfield type Fuel efficiency k, g/pass km (g/t km) Armament Calculated g-load nmax (nA) Airplane cost

940/10700 261 (285) 13750 7170 13430 3020

900/10650

890/10050

870/12000

900/13100

12500 12400 13500 3000

11800 8980 11900 2250

13100 7600 10400 2500

13115 15000 15200 2050

416/524/660

295/335/440

295/335

436

858/525

concrete

concrete

concrete

concrete

concrete

267 mil $

327,4 mil $

1.1 Scientific and technological forecasting Statistics on the aircraft-prototypes, released earlier, provides statistical graphs of their parameters on various factors of interest. We construct a statistical chart and trend line depending on the time t extension: As the trend function it seems logical to take a linear dependence of parameter on time: X(ti )= a + bt. The unknown parameters of the trend: a = (BC нAD) (nC нA2); b = (nD нAB) (nC нA2); x 0 = a. Where = ; = ; = D= ,

B=6,8+7,73+8+7,5+7,66+10,06+10,06+9,22+10,32+7,54=84,89 A=1981+1983+1986+1988+1989+1991+1992+1997+2001+2005=19913 11

=1981*2+1983*2+1986*2+1988*2+1989*2+1991*2+1992*2+1997*2+2001 *2+2005*2==29909 D=1981*6,8+1983*7,73+1986*8+1988*7,5+1989*7,66+1991*10,06+1992* 10,06+1997*9,22+2001*10,32+2005*7,54=189732,79 a = (BC нAD) (nC нA2)=84.89*39826-19913*189732,79/10*39826199132=8,53 b = (nD нAB) (nC нA2) =10*189732,79-19913*84.89/10*39826-199132= 0,0005 The predicted value x = a + bt; =8,55+0,0005*2015=9,6.
11 10,5 10 9,5 9 8,5 8 7,5 7 6,5 6 y = 0,076x - 144,4

1980

1985

1990

1995

2000

2005

2010

2015

t

We construct a statistical chart and trend line which shows the dependence of the sweep angle on the cruiser speed. It is logical to take a linear dependence of parameter on time as the function trend: X(ti )= a + bt. The unknown parameters of the trend: a = BC нAD nC нA2; b = nD нAB nC нA2; x 0 = a. 12

Where = = ; D= = , ;

=908+847+846+932+920+892+882+864+869+900=8860 =35+28+35+35+37,5+30+30+30+30+35=325,5 =908*2+847*2+846*2+932*2+920*2+892*2+882*2+864*2+869*2+900*2=17720 D=908*35+847*28+846*35+932*35+920*37,5+892*30+882*30+864*30+ +869*30+900*35=288936 2 a = BC нAD nC нA =325,5*17720-8860*288936/10*17720-88602=29 b = nD нAB nC нA2=10*288936-8860*325,5/10*17720-88602= 4,25 The predicted value
x = a + bt.

40 38 36 34 y = 0,068x - 28,36 32 30 28 26 840 850 860 870 880 890 900 910 920 930 940

V

1.2 Analysis of the project situation Let's note the specific features of the development and the level of excellence achieved by aircraft of this type. 1. The characteristics of the propulsion system: Maximum thrust: A-380 engines with P = 31 100 daN; Specific fuel consumption: C = 0,38; The thrust-to-weight ratio: m = 0,176. 13

2. Weights characteristics: Full load ratio: R = 0,498; Load ratio of payload: K = 0,15. 3. Geometrical characteristics: Aspect ratio 7,54;10,32 Taper ratio 3,66;5 Sweep angle 30 ;37,5 Fuselage aspect ratio 6,92;9,95 Relative base of the chassis b 0,36;0,416 Relative track of the undercarriage b 0,171;0,29 4. Flying characteristics: Cruising Speed: v 900 km / hour. Let's us consider the basic ways and means to ensure the technical excellence of aircraft: - The application of supercritical airfoils, allowing to increase the relative thickness of the wing, as well as its span, with no significant increase in weight; - The use of vertical wingtips, weakening the cross -flow on the wing, and thus able to reduce vortex drag and fuel consumption. The numerical value of the effect of their use in combination with supercritical airfoils, may be 5%; - Extensive use of composite materials in the construction of load bearing elements (carbon composites), fairings and flaps (GRP). The effect of their use can be expressed in the reduction of the mass construction of aircraft by 10%. - The use of electric remote control, display and new equipment, which can lead to a decrease in its weight by 8%. - Installation of efficient modern turbofan engines.

14

2 DEVELOPMENT OF INITIAL SPECIFICATION Initial specification of the projected plane define the basic goals and objectives of its establishment, terms of use, asking a suitable values of key parameters and characteristics of the aircraft, scheduled to the conditions of its production and operation. 2.1 Functional Requirements Appointment of an airplane: the passenger long-range wide-body aircraft The main tasks performed by the base plane: transport of passengers over distances up to 8,200 km. The use cases and possible modifications: the plane can be used as a passenger (the number of passengers up to 420 people at 3 class version), as well as military transport (transport of equipment or landing). The task forces: passengers Crew of two people. Terms of basing class airport, the type of runway, the possibility of short takeoff and landing: plane based on the ground not less than Class I with a length of 3,000 m runway. 2.2 General technical requirements These requirements define the basic performance of the future aircraft, its reliability and safety. These include such requirements as: 1. Ease of boarding and landing 2. High cruising speed 3. High fuel efficiency 4. Good takeoff and landing performance 5. Ease of maintenance and repair 6. Cabin volume of passengers We make the ranking of claims. 1 X 0 2 2 1 1 2 2 X 2 2 0 2 3 0 0 X 1 0 0 4 0 0 1 X 0 0 Table 2 н Ranking of claims 5 6 1 0 3 1 0 1 2 2 9 1 2 8 X 1 2 1 X 4 15 4 6 1 2 5 3

1 2 3 4 5 6

The results of paired comparisons allow us to write the basic requirements for this aircraft in descending order of importance: 1. High fuel efficiency 2. Good takeoff and landing performance 3. Cabin volume of passengers 4. Ease of boarding and landing 5. Ease of maintenance and repair 6. High cruising speed 2.3 Aircraft performance requirements ╖ cruising speed

V

of 900 km / h (

= 0.83);

╖ cruising altitude

of at least 10 000 m;

╖ the length of the takeoff more than 3000 m; ╖ landing approach speed

l

at maximum takeoff weight is not

V

. is not exceeding 270 km / hour.

2.4 Production and technological requirements 1 2 3 steel and . Batch production. . Broad use of one-piece panels . The main construction materials are aluminum and titanium alloys, composite materials.

2.5 Performance Requirements 1 2 3 4 5 aircraft. . . . . . Ensurance of the availability of facilities for servicing, repair. Allow automated control of the major aircraft systems. Easily removable assemblies and components. Ensurance interchangeability of units of aircraft. Ensurance unification and standardization of products serviced

2.6 The technical and economic requirements 1. Usage of load-bearing and secondary structures made from advanced materials to reduce weight and, consequently, the cost of transportation. 16

2. The coefficient of fuel efficiency with the progress in engine and aerodynamic characteristics of high radically new airframe to 100g/tkm. We define a trend line V(Lp)

=13430+13500+1190010400+15200+7820+13000+9700+8170+9700=112820 =940+900+890+870+900+908+932+846+847+864=8897 =2=112820=225640 D=13430*940+13500*900+11900*890+10400*870+15200*900+7820*908+ +13000*932+9700*846+8170*847+9700*864=100816750 a=(8897*225640-112820*100816750)/10*225640-1128202=889 b=(10*100816750-112820*8897)/10*225640-1128202=0,0003

17

3 BASELINE CONFIGURATION DEVELOPMENT 3.1 Selection of the wing parameters Based on the analysis of aircraft we perform a selection of the optimal geometrical parameters for the designed wing aircraft. With the increase of the wing aspect ratio increases aerodynamic efficiency (K), hence, decreases G and, therefore, reduced transportation costs. It is also improved by increasing of the aircraft landing, which reduces the required length. On the other hand, it increases with the increasing weight of the wing, and, consequently, the weight of the payload.

Aspect ratio are assigned, based on these reasons: 8,5 High value of taper ratio lead to the strong tendency of flow separation and increase in induced drag at the wing tips. We choose from these considerations, and analyzing the statistics 4 . The designed aircraft wing is swept. Based on the foregoing analysis and statistics ( 30...37,5 ) take 32 . The relative thickness of the wing affects the weight of the wing and at cruising speed. With the increase of c the weight of the wing is reduced, but you have to reduce the cruising speed of the aircraft. In this regard, the relative thickness of the wing is assigned: 0 0,12 . In the low-wing, when placing the propulsion system on the wing have to do V-wing positive, in order to prevent the possibility of contact between power plant and ground during landing. Accepted V - wing is 60. 3.2 Selection of the fuselage parameters The round shape is preferred for pressurized fuselage, loaded by internal pressure, as it excludes the appearance of large local stresses in the shell and, therefore, provides the least weight design. Given the statistics ( 6,08...11,28 ) accept 10 . We select 1,2 and 2,8 . We will take the fuselage diameter, based on statistical data. Since this class of airplanes in the passenger cabin perform mainly three rows of armchairs with three chairs in a row, then using statistical data, the diameter of the fuselage is adopt d 6,2m .

18

3.3 The choice of the tail parameters Symmetrical airfoils with a relative thickness slightly less than it is for the airfoils of the wing, a slight aspect ratio and greater sweep angle are chosen for the tail. According to statistics accept: - For the horizontal tail aspect ratio 4 , taper ratio 3 , relative thickness 0,09 and the angle of sweep 35 . - For the vertical tail aspect ratio 1,5 , taper ratio 3.4 , relative thickness 0,09 and the angle of sweep 450 . The relative area of the horizontal and vertical tail is S 0,25; S 0,2 Volume coefficients 0,7 ; 0,085 . 3.4 Selection of the control surfaces parameters According to the table [1, Table 3.4] the high-lift devices selects to increase the lift of the wing slat and choose the double -slit sliding flap with a relative chord b 0,3 , angle 40 0 . It gives a maximum lift coefficient max 2,8 at the landing angle of attack 130 . Basic controls: controlled stabilizer, rudder, ailerons. The relative areas of the rudders: S 0,35 bB 0,25 0,3 b 0,1 ;
S

0,35 0,45 b 0,25 0,3 b 0,2 0,25 .

Angles of deflection:

25 30 15 20

20 25 20 25

15 18 15 18

3.5. Selection of the undercarriage parameters The main geometrical characteristics of the undercarriage with nose wheel are chosen from statistical data and guided by the recommendations of [1]: - Aircraft ground angle 0 0 - The angle of the wing root section 2 0 - Landing angle of attack 7 - Overturning angle of the aircraft 3 AK 7 0 2 0 0 0 50 - Landing gear offset angle 2 0 7 0 - Wheel base b 0,35...0,4 l 19

3.6 The relative position of the aircraft units The wing-fuselage: We choose the type of wing attachment scheme н low-wing. The angle of the wing 2 0 is chosen from the cruise flight conditions with minimal drag. The wing-feathers: Our plane is of the normal scheme. According to statistics, we choose the relative position of horizontal and vertical tail o LB 0,45 , L 0,3 . The volume coefficient is 0,7 ; 0,085 . 3.7 Selection of the propulsion system We select turbo-jet turbofan engine as the engines - prototype. On passenger aircraft according to airworthiness requirements should be at least two engines, which is dictated by the terms of take -off with a failed engine. According to the statistical data we accepted: at takeoff (N = 0, V = 0) 0,595 / . ╖ Fuel consumption - no more than 0,6 kg / h daN ╖ ╖ The thrust-to-weight of the engine - no more than 0,18. 3.8 Selecting the number of engines and their placement on the aircraft When we are landing with a roll (up to 4║), engines should not touch the ground, so we are placing the engines on pylons to creat e a large dihedral angle equal to 6║. Thus, we are taking at least two engines, and place them under an airplane wing.

20

4 DETERMINATION OF INITIAL AIRCRAFT PARAMETERS The initial data for further development will be the range 8200 km and passenger capacity н 420 seats. 4.1 Determination of the specific load on the wing The selected value of the specific load on the wing is tested on the following conditions: - Providing a given speed approach:

p

0

C

y max

V

2 .

30,2(1 m )
2

Then
p0 2,8 71,52 677,1 daN m 30,2(1 0,3)

- Provision of a given cruising speed on the estimated altitude:
0,297 2502 p0 C xa 0 0,029 7,02 785,7 daN m 13(1 0,6m ) 13(1 0,6 0,3)
''

V

2

2

Accepted minimum value p0=677,1 daN/2. 4.2 Aerodynamic parameters

The coefficient of the lift slope in the subsonic zone:
D0 k

Then
D0 1,02 0,046 7.02

0,85

max

- aerodynamic efficiency in cruise mode; - the maximum aerodynamic efficiency;

max

max

1 2 C
X
0

D0

1 13,9 ; 2 0,0286 0,046

0,85 14,677 11,8 .

21

5 IDENTIFICATION OF AIRCRAFT THRUST NEEDS Aircraft 0 100 / m0 g , where P0 - the total takeoff thrust of all engines daN. 1. Flying at a cruising speed V at altitude : 2. Provision of a given length runway : =1,05[ + (+
1 1,2 677,1 1 0 1,05 (0,02 10 2 2200 2

)]

) 0,256

3. Taking off with an engine failure on takeoff: = ( + );
1.5 4 1 ( 0.03) 0,245 4 1 10.8 Select the maximum thrust-to-weight ratio 0 0,256 . 0

22

6 DETERMINATION OF AIRCRAFT TAKEOFF WEIGHT 6.1 Definition of the initial take-off weight
m0

m m 1

57330 225 411107kg 1 0,27 0,1 0,4 0,09

This is take-off weight in the zeroth approximation, determined by statistical data [1, c.130]:
m 0,27; m 0,1; m 0,1; m

0,4; m 0,09.

6.2 Definition of the payload mass The commercial load refers to the target load for passenger aircraft, which includes passengers, baggage, cargo and mail. m = 1,3( m+q) n=1,3(75+30)420=57330 kg. 6.3 Definition of the equipment mass and service load Approximately the absolute mass of this group can be defined as the sum of the masses of the crew and equipment: m=m + m, Where m = m1 ╖ n=75*3=225 kg. m1 = 75 kg н average weight of one crew member for civil aircraft; n н number of crew members [1, . 215] take 3 people. The mass of equipment can be taken in a relative form and include the mass of equipment:
m m1 n m 0

75╖2 0,02╖411107 8372,14kg

6.4 Determination of the relative weight of construction We can take advantage of the statistical formula to determine this mass: m0 5.5 1 1 ╖ ╖ 2 0.065, m ╖ ╖n A 1000 p0 p0 So,
250239,1 8,5 5,5 1 0,065╖10╖1,2 0,15 0,08 0,212 m 0,03╖0,664╖2 1000 677,1 677,1 [1 c.130]; m 0,25.

23

6.5 Determination of relative weight of the fuel system The relative weight of the fuel system is given by: k , where k 1,02...1,08 - coefficient determining the fraction of the mass of pipes and other equipment included in the fuel system to its total mass. We assume in the calculations k ╨▀ 1,02 . The relative mass of the fuel consists of the following components. Fuel weight for aircraft with a long cruising flight phase can be written as + + . Fuel weight for cruise flight without fuel burning ,
8200 440 0,494 0 0,391 900 70 11,8

Given the impact of burnout on the range =
m

0,391 0,314 1 0,625 0,391

For takeoff and landing = (1 н 0,03m)
m

.

(1 0,03 6)

0,0055 11 0,05 1 0,004 11

Aeronautical margin =
m

.

0,9 0,494 0,03 13,9

Other fuel Complete relative weight of fuel will be equal to:
0,314 0,05 0,03 0,006 0,4 .

As a result, we can determine the relative mass of the entire fuel system: 1,02 0,4 0,408 .

24

6.6 Determination of the relative masses of powerplant Knowing the needed thrust-to-weight ratio (power available is possible to determine the relative weight of the powerplant = k - for turbofan engine;
1,78 0,13 0,256 0,08

, it

6.7 Determination of relative weight of the equipment and controls You can use the following statistical relationships to determine this weight. Passenger magisterial aircraft with m0> 10000 kg: + 0,06 + , where

Then

= 250239,1 kg; n=420.

250 30 420 0,06 0,02 0,109 250239,1

=0,11 6.8 Definition of take-off weight of the first approximation Determine take-off mass of the aircraft of the first approximation:
I m0

1 .

57330 225 8372,14 412044,6kg 1 0.25 0,08 0,4 0,11

25

7 DETERMINATION OF THE ABSOLUTE DIMENSIONS OF THE AIRCRAFT 7.1 Selection of engines According to the required thrust-to-weight ratio P0 of engines we can find the total thrust for take-off mass m10:

P0

1 g m0 9,81 412044,6 P0 0,256 103479,23daN 10 10

and we find the thrust of one engine
P0

P0

n

25869,8daN

where n - number of engines on aircraft. Engine PW4062.P=28600 daN 7.2 Determination of the fuel mass and volume Required mass of fuel is the following
mT
1 mTC m0 0,41 412044,6 161586,12kg KTC 1,02

The volume of the fuel
UT mT

T

161586,12 201,98m 800

3

Volume of the fuel tanks U = U +U=201,98+35,8=237,8 m3 where U = = 7.3 Specifying the wing parameters We determine the wing loading 0=677,1 daN/m2, corresponding winf area for takeoff weight of the aircraft m10 = 412044,6 kg is:
1 m0 g 412044,6 9,81 S 596,9m p0 677,1 2

Wingspan l S 8,5 596,9 71,23m
2S 1 l Central chord b0 2 1

Tip chord bk

2 596,9 3,35m 1 4 71,23
S 2 4 596,9 13,4m l 1 4 71,23

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Mean aerodynamic chord bA b0 13,4 8,93m Sweep angle =32, dihedral angle must be given the choice of scheme the plane.V = 6. We find the relative sizes and the chord along the span ailerons, spoilers, absorbers of the uplifting force, the flaps, and slats [1, 394]. We select the shape, size and location of the terminal lenses of vortices according to the given statistics. 7.4 Definition of the plumage parameters We define the relative position of horizontal and vertical tail,using the volume coefficients and :
LO LBO ba S 8,93 596,9 AO 0,7 25 m; S O 149,22 lS 71,23 596,9 ABO 0,078 27,7 m. S BO 119,38

2 3

2 3

The scale and the chords tail are defined the same way as similar to the size of the wing relative to the selected parameters ,, , : l O O S O 4 149,22 24,43 m; l BO BO S BO 1,5 119,38 13,38 m;
b bkBO 2 1 2 1 2 1 2 1

b0
b0
b

O

BO

S O l O S BO l BO S TO lTO S BO l BO
O

2 149,22 1 3 24,43 2 119,38 1 3,4 13,38 2 3 149,22 1 3 24,43 2 3,4 119,38 1 3,4 13,38

3,05 m, 4,05 m,

9,16 m,
13,79 m,

A O

b

ABO

2 b0 3 2 b0 3

BO

2 9,16 6,1 m, 3 2 13,79 9,19 m. 3

7.5 Determination of the fuselage dimensions We apply a two-class layout: II and III classes, the fuselage's diameter is equal to 6 m. We find the length of the fuselage, the length of fore and aft of its parts L D 11 6,22 68,42; D 1520 3 2 650 2 50 2 130 6,220 m, LH .. H .. D 1,1 6,22 6,842 m, L.. .. D 2,2 6,22 13,68 m. 27

7.6 Determination of the undercarriage parameters Basic parameters for the adopted the scheme of chassis are determined by the following: noc 3 AK 7 0 2 0 0 0 50 , where - Landing angle of attack; н twist angle of the root section; - aircraft ground angle; - Landing gear offset angle 2 0 7 0 ; - Wheel base b 0,38 l 25m ; - longitudinal position of nose and main wheels b a 25 23,5 1,5 m; a 0,94 b 24,53 m; - Wheel track =11m. The aircraft is designed for operation from airfields in class A (runway length> 2550 m, Rekv<450 kN). We select wheels on the parking load: on the main strut - CT 86; on bow strut - 88 CT.
m g a 412044,6 9,81 23,5 237476075N n ae 16 23,5 1,5
m g a 294317,6 9,81 0,94 169626,26 N n ae 16

= 240 kN.

P

P

m

m 412044,6 294317,6 N 1,4 1,4

Pick up the wheel 1500*500: P 260000N P 180000N

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8 AIRCRAFT WEIGHT ESTIMATION Masses of the main aircraft assemblies and parts are determined during the weight estimations, a list of equipment is estimated in groups with their weight , and weight of the target and service loads specified. Further refinement of the take-off weight of the aircraft is the result of the calculation. Weight summary of the aircraft, which is based on the weight calculation, determines the aircraft take-off mass of the second approximation. 8.1 Wight summary The coefficient of the mass aircraft return at full load
m

m mc m

125729,214 32963056 37084,014 195776,79

m0 m m 100% mo m0
k

407460,7 195776,99 51,95% 407460,7

The coefficient of mass impact on payload
m 57330 100% 100 14,07% m0 407460,7

NAME 1. Construction Wing Fuselage Plumage Landing gear 2. Propulsion syst Engines Eggregates 3. Equipment and 4. Empty aircraft 5. Equipment and load Crew Office equipment 6. Zero fuel 7. Payload Passengers Luggage, mail 8. Fuel 9. Full load 10. Take-off weight

Table 3 н Wight summary of aircraft mi, kg 125729,21 52741,7 37084,014 23871,8 12031,7 em 32963,56 32317,22 646,3 control 37084,014 195776,79 service 8597,14 225 8372,14 341533,56 57330 31500 12600+2100 162984,28 407460,7

mi 0,308 0,129 0,091 0,058 0,029 0,08 0,079 0,0015 0,091 0,480

0,14 0,4

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9. AICRAFT LAYOUT 9.1 Aircraft volume-weight assembly Placing a full load of equipment and the aircraft must meet the following requirements: - Ensuring the best conditions for the crew; - Creating a comfortable environment for passengers; - ensuring maximum efficiency of equipment and systems; - Rational use of the internal volume of the fuselage and wings; - Achieve the desired alignment for all possible options for loading the aircraft, which is achieved by placing a variable and consumed load (payload, fuel) as close as possible to the center of mass of the aircraft or symmetri cally with respect to it. The layout of toilet facilities: If duration is more than 4 hours, and there are more than 200 passengers we take a toilet for 50 passengers. We get 420/50 = 8 toilets. The area of the toilet 1.05 * 1.05 m The wardrobe layout: two wardrobe place in the nose and the tail of the fuselage. S=0.035*420=14.7m2 V=0.05*420=21m3 We will place the 14 stewards on a guide wire for 30 passengers inside. Installation of engines: Setting of engines in the modern aircraft must meet the following requirements: - Ensure a minimum increment of weight and drag of aircraft; - Ensure easy mounting and dismounting of engines, as well as easy access to all units during the inspection; - Allow for rapid containment and extinguishing a fire in the engine. The design of engine mount must provide a compensation for thermal deformations and damping engine. 9.2 The structural layout of the aircraft We are choosing construction materials: D16AT material is used for making spar caps, machined panels. OT4-1 material is used for the manufacture of skin and highly loaded parts. The material VT22 is used for the manufacture of parts of the undercarriage, connecting rods, bellcranks. The material 30KhGSA for the manufacture of highly loaded welded assemblies, hot-parts, power frames, brackets. Material of the steel VNS-5 is used fo under engines frames. It has high impact strength. 30

Material Steel CH-4 is used for production of the elements of honeycomb panels made of thin sheets. Well sealed, stamped and soldered. Composite materials are used for the manufacturing of slats, spoilers. Thick covering, supported by a set of stringers, semi -monocoque fuselage perceives design (stringer), the power factors. Power frames for uniform load distribution are set to the places of focused approach forces. Ordinary frames have no virtually load and they are intended primarily to maintain the shape of fuselage. Wing three rails has a bends construction, so strut is required, as in the caisson to make cuts impossible, to secure the main landing gear. This constructive - power circuit design provides the least weight with good strength characteristics. During the designing of power circuit we are guided by general principles of obtaining power structures of minimum weight: - Transfer of effort on the shortest path; - Maximum use of the height of the building, working on a bend; - The use of thin-walled closed-loop transfer torque; - The combination and integration of force elements to transfer loads acting at different times and under different loading cases; - Minimum disturbance smoother flow of power of various concentrators (notches, holes, sharp corners, sharp changes in cross section), leading to weight reduction and design resource.

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10 AIRCRAFT BALANCE 10.1 Choice of balance range The valid range of center of gravity position depends on the type of aircraft and, first of all, the shape of the wing i scheme, as well as the parameters of longitudinal control. Bounds of the allowed range are usually determined by the calculation of the longitudinal stability and controllability of the aircraft. In the early stages of design, when these calculations are still not allowed, balance range is chosen approximately, based on statistical information. The initial or baseline center of gravity of fully loaded aircraft (m0) must lie roughly in the middle of the acceptable range 0 35 , m 0,26 0,3 0,26 н 0,30 - for aircraft with swept wings. We have m X m X a 39119 36819 100% 0,309
ba 8930

Table 4 - Center of gravity positions table
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 Propulsion systems (internal) Propulsion systems (outdoor) 3rd class passengers (first salon) 3rd class passengers (second salon) 3rd class passengers 18324,8 18324,8 11212,5 20475 9262,5 Name of aggregates, cargoes belonging to the point Fuselage (front compartment) Fuselage (cylindrical part) Fuselage (fin compartment) Wing Horizontal tail assembly Vertical tail assembly Nose landing gear Main landing gear (1) Main landing gear (2) m, kg 7537 37623,52 8114,4 52993 8560,52 5707,02 4076,44 8152,9 8152,9 x, m 5,115 34,65 60,246 39,999 63,03 63,45 7,58 5,625 40,604 40,604 41,781 39,002 30,074 37,588 16,719 35,635 51,468 mx, kg*m 38551,755 1303308,47 489006,202 2119667,01 539569,576 362110,419 30899,4152 22929,975 331040,352 331040,352 341370,076 317979,406 551100,035 688792,582 187461,788 729626,625 476722,35 y, m 5,955 5,955 6,321 5,865 7,809 12,935 1,504 3,775 1,564 3,721 1,564 4,187 2,95 3,721 6,168 6,168 6,168 my, kg*m 44882,835 223988,51 51291,122 310803,95 66849,101 73820,304 6130,9658 15388,561 12751,136 30336,941 12751,136 34136,192 54058,16 68186,581 69158,7 126289,8 57131,1

32

(third salon) 18 19 Aircrew Payload, mail and passengers' baggage (in front of the center section) Payload, mail and passengers' baggage (behind the center section) Fuel (in front of the center of gravity) Fuel (with the center of gravity) Kitchen 225 7020 4,042 16,743 909,45 117535,86 6,322 4,392 1422,45 30831,84

20

11360

46,249

525388,64

4,392

49893,12

21 22 23

64634,4 96951,66 8350

36,937 43,949 23,546

2387400,83 4260928,51 196609,1

5,865 6,186 4,524

379080,76 599742,97 37775,4

The coordinates of the aircraft center of gravity for all possible options to operate the aircraft loading are defined
m

mii

mi

ym

miyi

mi

5.74m

The relative distance of the center of mass of the aircraft is then converted to the relative position along longitudinal axis

m

m X ba

where XA - coordinate along the X axis. 10.2 Specified loading conditions 1. Takeoff weight of aircraft released landing gear: =

m x m
i

ii

16018908,42 39,29 m;=0,309. 407644

2. Takeoff weight of aircraft retracted landing gear: m i x i = 16008578,69 39,27 m, =0,306. = 407644 mi 3. Landing weight, released landing gear: m i x i = 11954589,01 39,10 m, =0,256. = 305733 mi 4. Landing weight, retracted landing gear: m i x i = 1194425,9 39,07 m, =0,251. = 305733 mi 33

5. Maximum range option (without payload with additional fuel capacity, released landing gear) m i x i = 14507561,8 39,51 m, =0,301. = m i 367201,92 6. Maximum range option (without payload with an additional supply of fuel, retracted landing gear) m i x i = 14497232,07 38,48 m, =0,298. = 367201,92 mi 7. Maximum range option (landing weight, released landing gear) m i x i = 13588954,96 39,62 m, =0,313. = 342964 mi 8. Maximum range option (landing weight, retracted landing gear) m i x i = 13578625,23 39,59 m, =0,310. = 342964 mi 9. The empty plane (in the parking lot, released landing gear) m i x i = 7071098,227 38,98 m, =0,271. = 181377,88 mi Maximum front center of gravity position н landing configuration, retracted landing gear (0,152) Maximum rear center of gravity position н maximum range configuration, landing weight, retracted landing gear (0,152) (0,346)

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11 DEVELOPMENT OF A GENERAL DRAWING VIEW AND AIRCRAFT TECHNICAL DESCRIPTION General information. Passenger aircraft to transport 420 passengers at a distance of 8200 kilometers was designed. The aircraft is made of a normal scheme, with a low-wing, double-deck, cargo hold and a kitchen on the lower deck. Aircraft structure. It consists of semi-monocoque fuselage, wing spars three, caisson-type chassis' beam. Slats and two slotted flaps are used to improve the landing characteristics. The nose strut is removed on the forward flight, the main - in the fuselage and wing. Coupled brake wheel are at the bow racks, four-wheeled trolley with brake wheels are on the major Control of the aircraft. Wire control system with quadruple redundancy is used at this aircraft. This reduces the weight of the aircraft and simplifies the automation process of piloting. Aircraft equipment and systems. Different equipment is used at this airplane: meteorological radar, navigation radar, laser locator station of range navigation, which provides automatic approach to a height of 5 meters; it is possible to install the runtime system of the fully automatic flight (or by the instructions of the software from the ground).Fire alarm system has four stages, triggered automatically or manually, fifth stage - only by hand. Defrosting system: nose of the wing and aircraft keel are heated by air, drawn from the engines, glass of the cockpit and the front edge of the stabilizer are heated by electricity. Aircraft is adapted for usage by state, so it installed sensors that measure the state of parts of the airframe and its engines. Fiber-optic cables that can significantly reduce the mass of the system and increase its reliability are used in the wiring. Propulsion system. The aircraft is fitted with four engines PW-4062. Engines are mounted by using pylons to power the wing ribs. Cruising speed 900km / h; Flight distance 8200km; Cruising altitude 10 km; Take-off weight 407460kg; The maximum payload 57330kg; Wing loading 677.1 daN/m2 Wing area 596.9m2; Engines Turbofan 4 * PW-4062; The total static thrust 633.23daN; The crew of 3 persons

35

CONCLUSION Aircraft for 420 passengers, a range of 8200km is designed according to the assignments. Its basic parameters were calculated: mass, capacity, etc. The aircraft layout and alignment of the aircraft was produced.

36

BIBLIOGRAPHY 1. , .. []: [ ]. - .: , 1991. н 400.: . 2. [] : . / .. ; . н , 1996. 76. 3. []: [ ] / [.. , .. , .. . .. ]. н 3- , . н .: , 1983. н 616 .

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