1. 1 Hydrogen Business Jet Preliminary Design Review Team III Derek Dalton Megan Darraugh Sara DaVia Beau Glim Seth Hahn Lauren Nordstrom Mark Weaver

2. 2 Design Requirements Alternative fuel: l H 2 Mid-sized –8 passengers Ultra-long-range business jet –Providing non-stop service between locations such as Los Angeles-Tokyo Range5,700nmi Passengers8 Cruise Speed0.80M

3. 3 Market Overview Projected 10-year revenue is $50B for the entire ultra long range market Acquire 15% market share within 10 years –Approximately 20 aircraft sold annually –$1.2B in potential annual sales, 2% of total business aviation market Expect to enter market in 2040 –Assuming $12B in development costs, will break even in 10 years

4. 4 Hydrogen Business Jet

5. 5 Front View 83.5 ft 28 ft 8 ft

6. 6 Side View 123 ft 26.45 ft 12 ft

7. 7 Compliance with Requirements RequiredDesign GTOW (lbs)£70,00058,700 V cruise (KTAS)³460460 H cruise (ft)³40,00040,000 Range (nmi)³5,7005,700 Landing Field Length (ft)£5,6005,370 Thrust per Engine* (lbs)£15,00012,630 Fuel Weight£11,70011,600 *Cruise Altitude Rated

8. 8 Carpet Plots L/D and Fuel Weight are constraining parameters GTOW58,700 lb AR11.9 T/W0.43 W/S100 lbs/ft 2 Design Point

9. 9 Carpet Plots Design Point

10. 10 Flight Envelope

11. 11 V-N Diagram Load Factors V stall 117 kts (SL) Altitude28,000 ft N lim 3.2g N ult 4.8g

12. 12 Mission Performance Distance (nm)Time (mn)Fuel (lbs) Speed (Mach)Altitude (ft)Thrust (lbs)L/D StartEndStartEnd Taxi Out- 1043 000 -- - Take Off-0.422 0 0 0 018,09116 Climb252.436.37670.3037,68618,8413,82216.5 Cruise5501.4719.495520.837,68640,0003,8223180 14.6 Hold 39.1488 -- ---- Descend127.127.11430.340,0000 --17 Taxi In- 1043 -00--- Reserves- -1052- - - -- - Flight Total5880.9782.812126 ------

13. 13 Twin Spool Turbofan Unmixed flow Bypass ratio: 4.5 Total pressure ratio: 25 Weight: 4,400 lbm Diameter: 3.6 ft Thrust (SLS): 33,800 lbf SFC (SLS): 0.14 lbm/lbf-hr

14. 14 Fan Study at SLS

15. 15 Structures & Materials Wing & fuselage skin: Carbon Epoxy laminate –Core in laminate adds stiffness for little additional weight –The laminate can be compared to an I-beam: Skins act as the I-beam flange Core materials act as the beam’s shear web Pylons: Titanium (Ti – 6Al- 4V) –Good for high load, poor shear properties

16. 16 Structures & Materials Lift was simplified to a point load at the aerodynamic center since the specific airfoil was not chosen Engine Weight Wing Weight Lift Force 11.5 ft 16.9 ft 21.5 ft Fuselage Ribs/stringers: –Al 2024 –Al 7075 –Al-Li alloy in future Landing gear: –Steel 300M Spar: –Aluminum 7175T66 will be used to get a 1.7 safety factor

17. 17 Landing Gear Tricycle configuration –Better visibility, good maneuverability –Requires proper balance to ensure braking and steering effectiveness Oleo-pneumatic shocks Diameter Rear Tire 29.2 in Width Rear Tire 9.2 in Diameter Front Tire 20 in Width Front Tire 6.6 in

18. 18 Supercritical Airfoils Relatively high cruise speed cause local shocks on most airfoils Supercritical airfoils reduce the severity of the shocks by distributing the pressure over the entire chord

19. 19 Airfoil Selection Unable to choose an airfoil because of limited data available on specific supercritical airfoils Most aircraft with transonic cruise have airfoils tailored to their specific mission

20. 20 Weight Breakdown Wing5,200 lb Fuselage18,260 lb Landing Gear2,545 lb Structure Total27,000 lb Engines4,380 lb (x2) Fuel11,600 lb Systems & Equipment10,160 lb Empty Weight43,570 lb Payload2,730 lb GTOW58,700 lb

21. 21 W landing gear, front W passengers W baggage W crew W fuel 2,3,4 A W fuel 1 W landing gear, main W verticle tail W horizontal tail W wing W engine W fuselage Weight Location X 123 ft 61.6 ft W fuel 2,3,4 CW fuel 2,3,4 B Neutral Point

22. 22 Fuel Storage 12 ft 78 ft 35 ft 8 ft Pax Area 3,4 2 1 Passengers 2 4 3 1 1)D = 8 ft, L = 43 ft, V = 2027 ft 3 8575 lb LH 2 2)Each Section: D = 3 ft, L = 25 ft, V = 169 ft^3717 lb LH 2 Tank 2: V = 509 ft 3 2152 lb LH 2 3)Each Section: D = 1.5 ft, L = 25 ft, V = 41 ft 3 172 lb LH 2 Tank 3: V = 122 ft 3 516 lb LH 2 4)Each Section: D = 1.5 ft, L = 25 ft, V = 41 ft 3 172 lb LH 2 Tank 4: V = 122 ft3516 lb LH 2 Total: V = 2780 ft 3 11760 lb LH 2 Nose:2*8 = 16 ft Tail:3.6*8 = 28 ft Total:78 + 16 + 24 = 123 ft Fuel Weight = 11760 lb LH 2 Density = 4.23 lb/ft 3 ABC

23. 23 C.G. Travel Neutral Point

24. 24 Dynamic Stability Vertical Tail Volume 1 Engine Out Takeoff 350 ft 2 Cross-wind Landing 307 ft 2

25. 25 Cost Acquisition Cost –Based off Average of FLOPS and Historical Trend Data –Both took into account increased technology as weighted factors Direct Operating Cost (DOC) –$5/gallon for Hydrogen –4 Flight Crew –Weighted Factors for Engine/Airframe Labor, Burden, and also Insurance –Approx. $50,000/departure at 200 departures per year

26. 26 HBJBBJG550 GTOW (lbs)58,700171,00091,000 Wing Area (ft 2 )58713451137 Span (ft)83.5311791.5 Λ 1/4c (Degree)303427 T/W0.430.320.352 W/S (lbf/ft 2 )10012780 V cruise (KTAS)460450460 H cruise (ft)40,00039,00040,000 Range (nmi)570062006500 AR11.9107.4 (L/D) max 1717.518.4 Landing Distance (ft)322425492767 Takeoff Field Length (ft)445956435934 Cost ($2006 Millions)605238 DOC ($2006)407029001820

27. 27 Outstanding Issues Stability Airfoil Selection and Aerodynamic Analysis Detailed Structural Analysis FAA Certification Research and Development Cost Analysis

28. 28 Questions

29. 29

30. 30 Hydrogen Safety Explosion Hazard –Leakage and Boil over Explosion –Similar to Jet Fuel Characteristics Proper Care and Materials Needed –Materials need to withstand very Low Temperatures –Safety Relief Valves, Purging, Sensors, and Sophisticated Seals Proven As Safe as Jet Fuel –No Detonation in Free Atmosphere –Tested and Comply with Present Regulations Fire Hazard –Boils off –No Fire Carpet –Fast Burn with Low Radiation.30 Caliber Armor PiercingPlaced in BonfireCharred Remains

31. 31 Hydrogen Fueled Engine Hydrogen has lower Flame Temperature –Reduced Turbine Inlet Temperature resulting in decrease in thrust Premixing almost necessary for proper combustion Other Slight Modifications needed

32. 32 Possible Fuel Cell APU Advantages –Reduces Size of Engine –Hydrogen Already onboard –Can be stored in empty wings –Reduced Noise Disadvantages –Needs Several Megawatts of Energy –Current APU’s producing only a few Megawatts and outweighing turbines

33. 33 Fuel Supply/Engine Modifications High Pressure Pump –Centrifugal Pump at approx. 150 RPM –Move LH2 from tanks to combustor Heat Exchanger –Transform Liquid to Gas before Combustion –Needs to increase temperature to about 150-300 K Purging System –Flush Air from Pipes Added Sensory –Sense Leakage Proper Materials –Perform at very low temperature

34. 34 Cryogenic Liquid Hydrogen Critical Temperature –-400 °F Critical Pressure –188 psia Cryogenic Storage –-423 °F –30 psia Requires 30% of Heating Value to Liquefy (15,000 BTU/lbm)

35. 35 Cryogenic Tank Current Cryogenic Tanks –Carbon Steel Alloy Outer Shell –Perlite Insulating Layer with Mylar wrapped Inner Shell –Al-Ni Inner Shell Future Cryogenic Tanks –Carbon Fiber Outer Shell –Graphite Fiber – Resin Matrix Composite Insulation –Advanced Composite Inner Shell LH 2

36. 36 Top View 8 ft

37. 37 FLOPS Input Moved Fuel from Wings to Fuselage Modified Heating Value to 54,000 BTU/lbm Added Composite Wing and Fraction of Structure Additional Weighted Factors for Fuel System to Include Cryogenics Increased Cost of Labor and Material

38. 38 Constraint: Landing

39. 39 Constraint: Landing

40. 40 Constraint: Landing

41. 41 Constraint: Landing

42. 42 V-N Diagram Support Gust Velocities –At V B, G = 60.4 ft/s –At V C, G = 45 ft/s –At V D, G = 22.5 ft/s n Gust = 1 + V G (K G GC Lalpha )/(498W/S) –K G =.88u/(5.3 + u) –u = 2W/S/(p*c bar *g*C Lalpha )

43. 43 Fan Study at 40,000 ft

44. 44 Engine Code Validation Bypass Ratio Pressure Ratio Thrust (SLS) Thrust (11 km) SFC (11 km) Cryoplane Engine 5.236.63129.5 kN24.77 kN 5.502 kg/(s-MN) HBJ Engine Code 5.236.6121.2 kN33.3 kN 6.2 kg/(s-MN)

45. 45 Landing Gear Size calculation –Used Business Twin equations –Table 11:1 – English Units Diameter or Width (inches) = A * W w B (where W w is the weight applied on each wheel)

46. 46 Structural Layout Guidelines –Never attach to skin alone –Structural members should not pass through cabins, air inlets, etc –Attach engine, landing gear, seats, etc to existing structural member –Design redundancy into structure –Mount control surfaces to spar Carry-through wing Added structural complexity with tanks above main cabin

47. 47 Spar Calculation

48. 48 Stability Formulas Add cg equ. Xbar=xi*Wi/Wi Static margin formula = Xn- X bar/C bar Equ for VHT

49. 49 ComponentC.G. Location (ft) Front Landing Gear15 Crew16 Baggage25.5 Furnishings33.5 Passengers40 Miscellaneous50 Fuel Tank 2,3,4 A29.5 Fuel Tank 2,3,4 B56.5 Fuel Tank 2,3,4 C83.5 Fuselage61.5 ComponentC.G. Location (ft) Fuel Tank 174.5 Main Landing Gear80 Wing58.9 Nacelle58.9 Engine58.9 Vertical Tail120 Horizontal Tail121 C.G. Location

50. 50 Stability Summary C.G. (ft)SMWeightnp-cg (ft) Wo 61.2516.84%40827.001.30 Wo+1/2 Fuel 61.4514.22%42019.281.10 Wo+1/2 Fuel+crew 60.5026.55%42919.282.05 Wo+fuel 61.6411.75%43211.550.91 Wo+fuel+all cargo 61.7410.57%57687.970.82 Wo+all cargo 59.3841.07%43399.003.18 ready for take-off 61.7410.57%57687.970.82 taxi & take-off 61.7310.67%57627.000.83 climb 61.6311.88%56944.210.92 cruise 60.1630.99%47378.552.40 decent 60.1231.40%47260.552.43

51. 51 Second Look at Stability W fuel 1 W fuel 5 Tank 5 = ¼ of original tank 1 Tank 1 = ¾ of original tank 1 Tank 5 = 1/3 of original Tank 1 Tank 1 = 2/3 of original Tank 1

52. 52 Additional Trade Studies Original1/41/3 Wo 61.561.3161.33 Wo+1/2 Fuel 62.4761.3761.09 Wo+1/2 Fuel+crew 62.0360.6360.35 Wo+fuel 63.9561.6261.10 Wo+fuel+all cargo 61.7360.2859.91 Wo+all cargo 62.4257.9360.05 ready for take-off 61.7360.2850.94 taxi & take-off 61.7260.2859.94 climb 61.6260.2659.95 cruise 60.1359.9960.07 decent 60.0959.9860.07 Original1/41/3 Wo 13.5416.1415.99 Wo+1/2 Fuel 1.0315.4519.00 Wo+1/2 Fuel+crew 6.6725.1128.68 Wo+fuel -18.1612.1718.89 Wo+fuel+all cargo 10.6129.6534.41 Wo+all cargo 1.6427.6332.59 ready for take-off 10.6129.6734.07 taxi & take-off 10.7129.6934.05 climb 12.0729.9533.95 cruise 31.3333.3932.31 decent 31.8433.5232.32 Center of Gravity (ft)Static Margin (%)

53. 53 Drag Polar

54. 54 1 Engine Out Calculations F v = q v *S v *C Fβv *  β v /  β*δR –q v (dynamic pressure at sea-level) = 48.61 slug/(ft*sec 2 ) –S v (vertical tail area) ft 2 –C Fβv (tail lateral lift force coefficient) = 0.55 –  β v /  β (free stream angle change) = 0.99 –δR (rudder deflection) = 0.35 radians T (thrust from 1 engine) = 12500 lb f M (total moment) = T*d E -F v *d V = 0 –Calculate with distances from center of gravity. Solve for needed vertical tail area.

55. 55 Cross-wind Landing C nβ = C nβw +C nβfus +C nβv -F pβ /(q*S w )*  β v /  β*(X cg -X p ) –Solve for C nβv when C nβ = 0 C nβv = C Fβv *  β v /  β*η v *S v /S w *(X acv -X cg ) –Solve for S v to find required vertical tail area

56. 56 Cost Support Historical Trend –$ = Aw e a M b R c A = 725.14 a = 0.1894 b = -.0519 c = 1.0777 50% Adjustment FLOPS –Composite Wing/Structure Factor –Increased Fuel System Factor –Advanced Technology Factor

57. 57 DOC Support

58. 58 DOC Support Contd

59. 59 DOC Support Contd

60. 60 DOC Support Cont