aircraft performance and design project code

1

Appendix A

Table of ContentsSpecifications ..................................................................................................................... 1Gross Weight Calculations ................................................................................................... 2Wing Loading .................................................................................................................... 4Cruise wing loading is least value ......................................................................................... 5Thrust to Weight ................................................................................................................ 6Mean aerodynamic chord ..................................................................................................... 71st CG Estimate ................................................................................................................. 82nd CG Estimate (Wing Addition) ........................................................................................ 8Tail Addition ..................................................................................................................... 92nd Wing Placement and Landing Gear Specifications ............................................................ 12Better Weight Estimates ..................................................................................................... 13Convergence of Weight Estimates ....................................................................................... 16Second Performance Analysis ............................................................................................. 18Time calculations .............................................................................................................. 23

Specificationsclc,clear

% Passengers% 10 family members (200 lbs.)% 2 pilots (180 lbs.)% 1 Flight attendant (180 lbs.)% Baggage per crew (20 lbs.)% Baggage per passenger (50 lbs.)

% Total Payload WeightWp = 10*200 + 20*3 + 10*50% Total Crew WeightWc = 3*180

% Cruise SpecificationsVcruise = .75 * 659.8 * 5280 / 3600; % ft/sHcruise = 40000; % Altitude (ft)% Stall SpecificationsVstall = 90 * 5280 / 3600; % ft/s

% Range Specifications (PHL to Bankok)R = 8721 * 5280; % ft% Loiter SpecificationsVloiter = .6 * 678.1 * 5280 / 3600; % ft/sHloiter = 30000; % ftE = 1800; % seconds

% Rate of Climb Specifications

Appendix A

2

RC = 1500 / 60; % ft/s

% POWER PLANT% More than 1 Turbofan

% MATERIAL SPECIFICATIONS% Composite Material

% Load Limit Factor Sepcificationn = 4.0;

% Take Off and Landing Distance Specification % s_T corresponds to Take-Off % s_L corresponds to LandingsgT = 6000; % ftsaT = 50; % ftsgL = 4000; % ftsaL = 50; % ft

% Densitiesrho40 = .00058727; % slugs/ft3rho30 = .00089068; % slugs/ft3rhoS = .0023769; % slugs/ft3

Wp =

2560

Wc =

540

Gross Weight Calculations% Jet specific fuel consumptions (sfc) with respect to the cruise and% loiter mission sections. % Values are derived form Table 3.3 (Raymer) for high-bypass turbofans % then converted to from lb/hr/lb to lb/s/lbCtcruise = .5/3600; % lb/sCtloiter = .4/3600; % lb/s

% Empy weight to gross weight ratio assumed from figure 8.1 (Anderson) for % an approximate gross weight estimate according to previously designed % aircraft depending upon their missions. NOTE: all ratios below are % derived from historical data except for empty-gross.WeW0 = .45;W1W0 = .97;W2W1 = .985;W5W4 = .995;

Appendix A

3

% B-52 Bomber specificaitons, desired LDmax is 21 while in the loiter % phase. The loiter phase will see the maximum lift to drag ratio while % the cruise mission segment will see a slight reduction in the LD % ratio according to page 22 (Raymer)LDmaxL = 21;LDmaxC = (.866).*LDmaxL;

% Cruise mission objective is for maximum range while loiter phase requires % flying for endurance or specified time.W3W2 = 1./(exp((R.*Ctcruise)./(Vcruise.*LDmaxC)))W4W3 = 1./(exp((E.*Ctloiter)./LDmaxL))

W5W0 = W1W0.*W2W1.*W3W2.*W4W3.*W5W4

WfW0 = 1.06.*(1 - W5W0)

denom = 1- WfW0 - WeW0;

% Gross, Fuel and Empty Weights, Fcap = Fuel CapacityW0 = (Wp + Wc)./denom %lbsWf = W0.*WfW0 %lbsWe = .45*W0 %lbs

Fcap = (Wf./5.64)*0.133681 %ft^3

W3W2 =

0.6160

W4W3 =

0.9905

W5W0 =

0.5800

WfW0 =

0.4451

W0 =

2.9566e+04

Wf =

1.3161e+04

Appendix A

4

We =

1.3305e+04

Fcap =

311.9510

Wing Loading% Wing loading calculations for various mission segments of the flight. The % The minimum wing loading will be selected.

% Take Off (Stall Velocity)CL_airfoil = 1.6;CL_highlift = .9;CL_max = (CL_airfoil+CL_highlift)*.9q = (0.5).*rhoS.*(Vstall.^2);

W1S = q.*CL_max

% Landingg = 32.2;Radius2 = ((1.23*Vstall)^2)/(.2*32.2);hf = Radius2*(1-cosd(3));sa2 = (50 -hf)/tand(3);sf2 = Radius2*sind(3);j = 1.15;N = 3;Ur = 0.4;sg2 = sgL - sa2 - sf2;

A = j.*N.*sqrt(2./(rhoS.*CL_max));B = (j.^2)./(g.*rhoS.*CL_max.*Ur);

C = B.^2;D = (A.^2) + (2.*sg2.*B);Z = sg2.^2;

x = [C -D Z];W2S = roots(x)

% CruiseCdo = .012; % Taken from B-52e = .6; % low wing from McCormickK = 1./(4*(LDmaxC.^2).*Cdo)W3S = (Vcruise.^2).*rho40./(2*sqrt((3.*K)./(Cdo)))

Appendix A

5

CL_max =

2.2500

W1S =

46.5920

W2S =

202.5173 115.6891

K =

0.0630

W3S =

38.9767

Cruise wing loading is least value% Wing loading value selected as minimum from cruise flight and used to % obtain wing area, and wing span thereafter. The aspect ratio is % derived from induced drag and oswald efficiency factor.WSmin = W3S

S = W0./WSminAR = 1./(pi.*e.*K)b = sqrt(AR.*S)

WSmin =

38.9767

S =

758.5505

AR =

8.4220

Appendix A

6

b =

79.9281

Thrust to Weight% Thrust to weight calculations for various mission segments. The highest % value is to be selected.% Take Offn = 4;CLadjust = .9*(1.7 + .5);Radius = (6.96*(Vstall^2))/32.2;ThetaOB = acosd(1-(50/Radius));Sa = Radius*sind(ThetaOB);y = (1.21*W3S)/(32.2*rhoS*CLadjust);TW1 = y/(sgT - Sa)

% Rate of ClimbTW2 = (RC/Vstall)+(.5*rhoS*(Vstall^2)*(W3S^-1)*Cdo)+((2*W3S*K)/(rhoS*(Vstall^2)))

%Thrust Matchinga = .267;c = .363;TW3 = a.*(.75).^c

%Transport StatisticalTW4 = .25

%Sustained TurnTWt = (.5*rho40*(Vcruise^2)*Cdo/W3S)+((W3S*K*(n^2))/(.5*rho40*(Vcruise^2)))

% CheckTWturnCheck = 2.*n.*sqrt(K.*Cdo)TWcruiseCheck = 2*sqrt(K.*Cdo)

TW = TW2

T = W0.*TW

TW1 =

0.0578

TW2 =

0.3143

TW3 =

Appendix A

7

0.2405

TW4 =

0.2500

TWt =

0.3016

TWturnCheck =

0.2199

TWcruiseCheck =

0.0550

TW =

0.3143

T =

9.2936e+03

Mean aerodynamic chord% cT = tip chord length, cR = root chord length, ybar = height of m.a.c., % cbar = spanwise location of m.a.c. and taper ratio lambda.cT = 7.2;cR = 12;lambda = 0.6;ybar = (b./6).*((1+2.*lambda)./(1+lambda))cbar = (2/3).*cR.*((1+lambda+(lambda.^2))./(1+lambda))

ybar =

18.3169

cbar =

9.8000

Appendix A

8

1st CG Estimate% Approximate locations and weights of all fuselage components. Used for % moment calculation. (Refer to the AutoCAD sketch for better % perspective).

% Enginex1 = 80.8115;w1 = 1.4*1644;% Flight Attendantx2 = 24.222;w2 = 180;% Bathroomx3 = 28.972;w3 = 130;% Fridge & Food/ Drinkx4 = 28.972;w4 = 400;% Passengersx5 = 56.157;w5 = (10*200);% Fuel Secondary and crew baggagex6 = 36.7970;w6 = 4767.5 + 20*3;% Electrical System and Fuel Pumpx7 = 43.2970;w7 = 200;% Pilotsx8 = 20.472;w8 = 2*180;%Passenger Baggagex9 = 70.342;w9 = 50*10;

% Total Moment aabout Aircraft NoseMn = (x1.*w1) + (x2.*w2) + (x3.*w3) + (x4.*w4) + (x5.*w5) + (x6.*w6)... + (x7.*w7) + (x8.*w8) + (x9.*w9);wn = w1 + w2 + w3 + w4 + w5 + w6 + w7 + w8 + w9;

% First center of gravity without wingsCG1 = Mn./wn

CG1 =

50.1750

2nd CG Estimate (Wing Addition)% Wing mean aerodynamic center located at fuselage CG1 aboveWwing = (2.5).*S;WwF = Wwing + (Wf - 4767.5)

Appendix A

9

% Distance from leading edge to aerodynamic centercbar1 = (.25).*cbar

% Distance between aerodynamic center and the center of gravitycbar2 = (.4).*cbar - cbar1

% Fuel weight had been incorporated in the wing weight value because of the % significant amount of weight added to the wings from the fuel. Process % follows that of Anderson.CG2 = (Mn + WwF.*(CG1 + cbar2))./(wn + WwF)

WwF =

1.0290e+04

cbar1 =

2.4500

cbar2 =

1.4700

CG2 =

50.8889

Tail Addition% Horizontal and Vertical tail volume ratios, values were averaged form data % ranges of aircraft with good stability characterisitcs.VHT = .5;VVT = .0425;

% Measuring from the nose of the aircraft, PL = Plane Length, PL2 = % approximate location of the mean aerodynmic center of the horizontal % tail.PL = 90;PL2 = (.965250965).*PL

% Locations of the horizontal and vertical m.a.c. from the center of % gravity of the aircraft (wings and fuselage). The horizontal is found % by subtracting the fuselage/wing combination center of gravity from % the overall length the Horizontal m.a.c. from the a/c nose.lHT = PL2 - CG2lVT = (.934033859).*lHT

Appendix A

10

% The planview area calculations for both horizontal and vertical sections.SHT = (VHT.*cbar.*S)./lHTSVT = (VVT.*b.*S)./lVT

% Assumed aspect ratio of the horizontal wing section ARH according to % Anderson. The horizontal span, root cord length and tip chord length % calculations are shown below respectively.ARH = 4;bt = sqrt(SHT.*ARH)crt = (2.*SHT)./((lambda + 1).*bt)ctt = lambda.*crt

% These are the coordinates for the location of the m.a.c. of the% horizontal tail. cHT is the spanwise distance from the right-most edge,% while yHT is distance away from the centerline. (Where the mirror of the% tail occurs)yHT = (bt./6).*(1 + 2.*lambda)./(1 + lambda)cHT = (2/3).*crt.*(1 + lambda + (lambda.^2))./(1+lambda)

% Vertical tail section aspect ratio, averaged from data range found on % page 441 (Anderson)ARV = 1.65;Lambda_Vtail = .85;

% In order below, the height of the vertical tail section, root chord % length, and tip chord length.hVT = sqrt(ARV.*SVT)crVT = 2.*SVT./((Lambda_Vtail+1).*hVT)ctVT = Lambda_Vtail.*crVT

% Similar to the horizontal tail section, cVT is the distance of the m.a.c. % measured spanwise from the right of the airfoil, while zVT is height % component of the m.a.c. location.zVT = (2.*hVT./6).*(1+(2.*Lambda_Vtail))./(1+Lambda_Vtail)cVT = (2/3).*crVT.*(1 + Lambda_Vtail + (Lambda_Vtail.^2))./(1 + Lambda_Vtail)

PL2 =

86.8726

lHT =

35.9837

lVT =

33.6100

SHT =

Appendix A

11

103.2940

SVT =

76.6663

bt =

20.3267

crt =

6.3521

ctt =

3.8113

yHT =

4.6582

cHT =

5.1876

hVT =

11.2472

crVT =

7.3692

ctVT =

6.2638

zVT =

5.4716

cVT =

Appendix A

12

6.8314

2nd Wing Placement and Landing Gear Specifi-cations

% Static margin given as 10% (Anderson), calculation of the aerodynamic % center of the wing body. Followed by calculation of the wing leading % edge location, Crlead, and wing center location, Xc. NOTE: that Xc % will be the location of the main landing gear for the aircraft. % Measurements are made from the nose of the a/c.SM = 0.1;Xn = SM.*cbar + CG2;Xacwing = Xn - VHT

Crlead = Xacwing - cbar1 - ((cR - cbar)./2);Xc = Crlead + cR./2

Xnose = (0.086872587).*PL

% Distance calculation of each wheel from the known center of gravity (as% opposed to the nose). Diagram can be seen on page 446 (Anderson). There% are two location for the focus of all aircraft weight, being the nose and% main landing gears. The main landing gear consists of a set of 2 wheels% (left and right) and therefore it is necessary to split the load at that% between between them.X3 = Xc - Xnose;X1 = CG2 - Xnose;X2 = X3 - X1;

Fm = (W0.*X1)./X3Fn = (W0.*X2)./X3

% Wheel Dimensions are calculated according to equation 8.82 (Anderson), % where AD is the A diameter coefficient and AW is the A width % coefficient, and so-on and so-forth.AD = 1.51;BD = 0.349;AW = 0.715;BW = 0.312;

% Main wheel diameter and widthMD = (AD.*((Fm./2).^BD))./12MW = (AW.*((Fm./2).^BW))./12

% Nosewheel diameter and widthND = (AD.*(Fn.^BD))./12NW = (AW.*(Fn.^BW))./12

Xacwing =

Appendix A

13

51.3689

Xc =

53.8189

Xnose =

7.8185

Fm =

2.7683e+04

Fn =

1.8832e+03

MD =

3.5081

MW =

1.1673

ND =

1.7488

NW =

0.6265

Better Weight Estimates%Front Cone AreaFront = pi*4*(4+sqrt(16.472^2+4^2))

%Main Cylinder AreaCenter = (2*pi*4*56.476)+(2*pi*4^2)

%Rear Elipse Area

Appendix A

14

Back = pi*8*6.25*17.052

%Total Wetted AreaWetted_area = Front+Center+Back

%Fuselage WeightW_Fuselage = Wetted_area*1.4

%Main Wing WeightW_Wings = 2.5*((7.2*36)+(36*2.4))

% Horizontal Tail WeightW_HStab = 2*((4.1192*9.9822)+(1.25555*9.9822)+(.5*.9398*6.6303))

% Vertical Tail WeightW_VStab = 2*((11.5656*5.2855)+(1.7618*11.5656))

% Landing Gear WeightW_LGear = .057*W0

% W_EngineW_Engine = 1.4*1644

% Other weightW_Other = (.1*W0)

% Total Empty WeightW_Empty = W_Fuselage+W_Wings+W_HStab+W_VStab+W_LGear+W_Engine+W_Other

Wf

% Take Off WeightW_TO = Wc+Wp+Wf+W_Empty

Front =

263.2745

Center =

1.5199e+03

Back =

2.6785e+03

Wetted_area =

4.4617e+03

Appendix A

15

W_Fuselage =

6.2464e+03

W_Wings =

864.0000

W_HStab =

113.5348

W_VStab =

163.0125

W_LGear =

1.6853e+03

W_Engine =

2.3016e+03

W_Other =

2.9566e+03

W_Empty =

1.4330e+04

Wf =

1.3161e+04

W_TO =

3.0592e+04

Appendix A

16

Convergence of Weight Estimates% Pluged the new W0 weight into the landing gear and other calculations and% redetermined We (still using WeW0 ratio from the beginning and Wf.% Repeated 16 times until all three weight values remianed steady.W_LGear2 = .057*W_TO;W_Other2 = (.1*W_TO);W_Empty2 = W_LGear2+W_Other2+W_Fuselage+W_Wings+W_HStab+W_VStab+W_Engine;Wf2 = WfW0*W_TO;W_TO2 = Wc+Wp+Wf2+W_Empty2;

W_LGear3 = .057*W_TO2;W_Other3 = (.1*W_TO2);W_Empty3 = W_LGear3+W_Other3+W_Fuselage+W_Wings+W_HStab+W_VStab+W_Engine;Wf3 = WfW0*W_TO2;W_TO3 = Wc+Wp+Wf3+W_Empty3;







Appendix A

17







W_LGear16 = .057*W_TO15;W_Other16 = (.1*W_TO15);W_Empty16 = W_LGear16+W_Other16+W_Fuselage+W_Wings+W_HStab+W_VStab+W_EngineWf16 = WfW0*W_TO15W_TO16 = Wc+Wp+Wf16+W_Empty16

W_Empty16 =

1.4735e+04

Wf16 =

1.4308e+04

W_TO16 =

Appendix A

18

3.2143e+04

Second Performance Analysis% 2nd analysis gross weight, empty weight, and fuel weight calculations.W0v2 = W_TO16;Wev2 = W_Empty16 ;Wfv2 = Wf16;

% Lowest wing loading value, as well as the wing loading specific to climb, % to be used in rate of climb calculations.WSn = W0v2./SWSclimb = (.97)*(.985).*WSn

% Rate of Climb % The data series/ arrays shown below are for air densities according % to their respective altitude, i.e. sea level, h = 0 and rho = % 2.3769*10^-3. The rate of climb graph is derived from this data % series being proportionally incorporated into thrut to weight % caluclations. According to Anderson, as altitude increases, Thrust % decreases proportional to rho^0.6 for turbofan engines.rho_series = [2.3769 2.3423 2.3081 2.2743 2.2409 2.2079 2.1752 2.1429 ... 2.1110 2.0794 2.0482 2.0174 1.9869 1.9567 1.9270 1.8975 1.8685 ... 1.8397 1.8113 1.7833 1.7556 1.7282 1.7011 1.6744 1.6480 1.6219 ... 1.5961 1.5707 1.5455 1.5207 1.4962 1.4719 1.4480 1.4244 1.4011 ... 1.3781 1.3553 1.3329 1.3107 1.2889 1.2673 1.2459 1.2249 1.2041 ... 1.1836 1.1634 1.1435 1.1238 1.1043 1.0852 1.0663 1.0476 1.0292 ... 1.0110 .99311 .97544 .95801 .94082 .92387 .90716 .89068 .87443 ... .85841 .84261 .82704 .81169 .79656 .78165 .76696 .75247 .73820 ... .72413 .71028 .69443 .67800 .66196 .64629 .63100 .61608 .60150 ... .58727 .41329 .39147 ] *(10^-3);

Altitude = [0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 ... 6000 6500 7000 7500 8000 8500 9000 9500 10000 10500 11000 11500 ... 12000 12500 13000 13500 14000 14500 15000 15500 16000 16500 ... 17000 17500 18000 18500 19000 19500 20000 20500 21000 21500 ... 22000 22500 23000 23500 24000 24500 25000 25500 26000 26500 ... 27000 27500 28000 28500 29000 29500 30000 30500 31000 31500 ... 32000 32500 33000 33500 34000 34500 35000 35500 36000 36500 ... 37000 37500 38000 38500 39000 39500 40000 47342 48475];

% Varying Thrust to Weight with increasing altitude. Note the sea level RC % is estimated to be in the upper 90's (ft/s), well over 5400 ft/min % which is 3.6 times shorter than desired specification.TWnew = (T./W0v2).*(rho_series./rhoS)

% Equations 5.116 and 5.113 (Anderson) to solve for RCmaximum with 2nd % analysis wing loading and thrust to weight specifications. The first % three lines below utilize the data series above to plot a graph of % the RCmax of the aircraft according to altitude. The following two % lines of code are used to calculate time to climb using the sea level

Appendix A

19

% values of the array. Linies 413 - 428 combine equations 5.116 and % 5.113 and then incorporate the change in thurst to weight with % respect to altitude change. The "combined" equation then set to a % a specific value of RCmax. According to Anderson, the absolute % and service ceilings occur when RCmax = 0, 100 ft/min respectively. % This translates to 0, and 5/3 ft/s. When plugged in and solved, x and % y yield the absolute and service ceiling densities. Looking in the % density tables in (Anderson), a corresponding height is found and % manually typed into Absolute_Ceiling and Servvice_Ceiling below. % (Values are interpolated by hand). These values were added to the % arrays above after the fact.Z = 1 + sqrt(1 + (3./((LDmaxL.^2).*(TWnew.^2))));RCmax = sqrt((WSclimb.*Z)./(3.*rho_series.*Cdo)).*(TWnew.^1.5).*... (1 - (Z./6) - (3./(2.*(TWnew.^2).*(LDmaxL.^2).*Z)));

% t_climb_min is converted to minutesRCmaxSea = RCmax(1,1)t_climb_min = (Altitude(1,81) - Altitude(1,1))./(RCmax(1,1));

syms x yAC = solve( sqrt((WSclimb.*(1 + sqrt(1 + (3./... ((LDmaxL.^2).*(((T./W0v2).*(x./rhoS)).^2))))))./(3.*x.*Cdo)).*... (((T./W0v2).*(x./rhoS)).^1.5).*(1 - ((1 + sqrt(1 + (3./... ((LDmaxL.^2).*(((T./W0v2).*(x./rhoS)).^2)))))./6) - (3./(2.*... (((T./W0v2).*(x./rhoS)).^2).*(LDmaxL.^2).*(1 + sqrt(1 + (3./... ((LDmaxL.^2).*(((T./W0v2).*(x./rhoS)).^2)))))))) == 0, x);

Absolute_Density = vpa(AC(1,:))Absolute_Ceiling = 48000 + (48500 - 48000)*((Absolute_Density - .00040045)/... (.00039099 - .00040045))

SC = solve( sqrt((WSclimb.*(1 + sqrt(1 + (3./... ((LDmaxL.^2).*(((T./W0v2).*(x./rhoS)).^2))))))./(3.*x.*Cdo)).*... (((T./W0v2).*(x./rhoS)).^1.5).*(1 - ((1 + sqrt(1 + (3./... ((LDmaxL.^2).*(((T./W0v2).*(x./rhoS)).^2)))))./6) - (3./(2.*... (((T./W0v2).*(x./rhoS)).^2).*(LDmaxL.^2).*(1 + sqrt(1 + (3./... ((LDmaxL.^2).*(((T./W0v2).*(x./rhoS)).^2)))))))) == 5/3 , x);

Service_Density = vpa(SC)Service_Ceiling = 47000 + (47500 - 47000)*((Service_Density - .00042008)/... (.00041015 - .00042008))

% Plotting and Marking Ceiling Dataint = 0:1:100;xx = Altitude(1,82).*(int./int);yy = Altitude(1,83).*(int./int);

plot(RCmax, Altitude,RCmax(1,83),Altitude(1,83),'rs',RCmax(1,82),... Altitude(1,82),'ks')hold onplot(xx,'k')plot(yy,'r')legend('RC vs. Altitude','Absolute Ceiling','Service Ceiling',... 'SC Marker','AC Marker')

Appendix A

20

grid onxlabel('RCmax (ft/s)')ylabel('Altitude (ft)')title('RCmax Vs. Altitude')axis([0 100 0 50000])

% 2nd Analysis Stall VelocityVstalln = sqrt((2./rhoS).*WSn./CLadjust)

% Landing Distance, where grv is gravity, AA is approach angle, Vf is flare % velocity, R is flight path radius,hF is flare height. Determined valu % is appr. 61% of the specified value 4000 ft. Well within constraints.grv = 32.2;AA = 3;Vf = (1.23).*Vstalln;Rv2 = (Vf.^2)./((0.2).*grv)

hF = Rv2.*(1 - cosd(AA))

SA = (50 - hF)./(tand(AA))SF = (Rv2.*sind(AA))SG = j.*N.*sqrt((2./rhoS).*WSn./CLadjust) + ... ((j.^2).*WSn./(grv.*rhoS.*CLadjust.*Ur))

SGn = SA + SF + SG

% 2nd Analysis Takeoff Distance, appr. 29% of specified value 6000 ft.SGT = (1.21).*WSn./(grv.*rhoS.*CLadjust.*TW(1,:))RR = (6.96).*(Vstalln.^2)./grvOB = acosd(1 - saT./RR)SAT = R.*sind(OB)TOD = SGT + SAT

WSn =

42.3747

WSclimb =

40.4869

TWnew =

Columns 1 through 7

0.2891 0.2849 0.2808 0.2766 0.2726 0.2686 0.2646

Columns 8 through 14

0.2607 0.2568 0.2529 0.2491 0.2454 0.2417 0.2380

Appendix A

21


0.2344 0.2308 0.2273 0.2238 0.2203 0.2169 0.2136


0.2102 0.2069 0.2037 0.2005 0.1973 0.1942 0.1911


0.1880 0.1850 0.1820 0.1790 0.1761 0.1733 0.1704


0.1676 0.1649 0.1621 0.1594 0.1568 0.1542 0.1516


0.1490 0.1465 0.1440 0.1415 0.1391 0.1367 0.1343


0.1320 0.1297 0.1274 0.1252 0.1230 0.1208 0.1187


0.1165 0.1144 0.1124 0.1103 0.1083 0.1064 0.1044


0.1025 0.1006 0.0987 0.0969 0.0951 0.0933 0.0915


0.0898 0.0881 0.0864 0.0845 0.0825 0.0805 0.0786


0.0768 0.0749 0.0732 0.0714 0.0503 0.0476

RCmaxSea =

97.7624

Absolute_Density = 0.00039147049586455985078970718599533 Absolute_Ceiling = 48474.603812655400139974897115356

Appendix A

22

Service_Density = 0.00041329397068659587824712505735577 Service_Ceiling = 47341.693318902524126111895057413

Vstalln =

134.1931

Rv2 =

4.2304e+03

hF =

5.7977

SA =

843.4309

SF =

221.4038

SG =

1.3875e+03

SGn =

2.4523e+03

SGT =

1.0764e+03

RR =

Appendix A

23

3.8924e+03

OB =

9.1935

SAT =

7.3569e+06

TOD =

7.3580e+06

Time calculations% Climb Segment

TCmin = (Hcruise - 0)./25ThetaC = asind(RCmax(1,1)./Vstalln)Rclimb = 39950./tand(ThetaC)

Appendix A

24

%Glide phase to loiter altitudeThetamin1 = atan(1./LDmaxL);Rglidemax1 = (Hcruise - Hloiter)./tan(Thetamin1)VLDmax1 = sqrt(2*WSn.*sqrt(K./Cdo)./rho40);VLDactual1 = sqrt(rho40./rho30).*VLDmax1;Vsink1 = VLDactual1.*sin(Thetamin1);

Tglide1 = (Hcruise - Hloiter)./Vsink1

% Glide Phase 2 to sea levelThetamin2 = atan(1./LDmaxL);Rglidemax2 = (Hloiter - 0)./tan(Thetamin2)VLDmax2 = sqrt(2*WSn.*sqrt(K./Cdo)./rho30);VLDactual2 = sqrt(rho30./rhoS).*VLDmax2;Vsink2 = VLDactual2.*sin(Thetamin2);

Tglide2 = (Hcruise - Hloiter)./Vsink2

T_Glide_Total = Tglide1 + Tglide2

% Loiter TimeTloiter = E

% Cruise TimeRnew = R - Rclimb - Rglidemax1 - Rglidemax2Tcruise = Rnew./Vcruise

% Cumulative trip timeT_trip_total = TCmin + + Tcruise + T_Glide_Total + TloiterThrs = T_trip_total./3600

TCmin =

1600

ThetaC =

46.7625

Rclimb =

3.7565e+04

Rglidemax1 =

210000

Tglide1 =

Appendix A

25

450.2751

Rglidemax2 =

630000

Tglide2 =

735.5675

T_Glide_Total =

1.1858e+03

Tloiter =

1800

Rnew =

4.5169e+07

Tcruise =

6.2236e+04

T_trip_total =

6.6821e+04

Thrs =

18.5615

Published with MATLAB® R2013a

aircraft performance and design project code

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