1

Hydrogen Business JetPreliminary Design Review

Team III

Derek DaltonMegan Darraugh

Sara DaViaBeau GlimSeth Hahn

Lauren NordstromMark Weaver

2

Design Requirements

• Alternative fuel: lH2

• Mid-sized – 8 passengers

• Ultra-long-range business jet – Providing non-stop

service between locations such as Los Angeles-Tokyo

Range 5,700nmi

Passengers 8

Cruise Speed 0.80M

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

Hydrogen Business Jet

5

Front View

83.5 ft

28 ft

8 ft

6

Side View

123 ft

26.45 ft

12 ft

7

Compliance with Requirements  Required Design

GTOW (lbs) 70,000 58,700

Vcruise (KTAS) 460 460

Hcruise (ft) 40,000 40,000

Range (nmi) 5,700 5,700

Landing Field Length (ft) 5,600 5,370

Thrust per Engine* (lbs) 15,000 12,630

Fuel Weight 11,700 11,600*Cruise Altitude Rated

8

Carpet Plots

L/D and Fuel Weight are constraining parameters

GTOW 58,700 lb

AR 11.9

T/W 0.43

W/S 100 lbs/ft2

Design Point

9

Carpet Plots

Design Point

10

Flight Envelope

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.60

1

2

3

4

5

6

x 104

Mach Number

Alt

itu

de

(ft

)

Thrust-Drag LimitStall LimitCeilingMaximum SpeedDesign Point

300 350 400 450 500 5500

1

2

3

4

5

6x 10

4 Placard Diagram

True Airspeed (knots)

Alt

itu

de

(fe

et)

Cruise

Dive

11

V-N Diagram

Load Factors

Vstall 117 kts (SL)

Altitude 28,000 ft

Nlim 3.2g

Nult 4.8g

12

Mission Performance

 Distance

(nm) Time (mn) Fuel (lbs)Speed (Mach) Altitude (ft) Thrust (lbs) L/D

          Start End Start End  

Taxi Out -  10 43  0 0 0  - -  - 

Take Off - 0.4 22 0  0  0  0 18,091 16

Climb 252.4 36.3 767 0.3 0 37,686 18,841 3,822 16.5

Cruise 5501.4 719.4 9552 0.8 37,686 40,000 3,822 3180  14.6

Hold   39.1 488  - -  - - - -

Descend 127.1 27.1 143 0.3 40,000 0  - - 17

Taxi In -  10 43  - 0 0 - - -

Reserves -   - 1052 -  -   -  - -  - 

Flight Total 5880.9 782.8 12126  - - - - - -

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

Fan Study at SLS

0 2 4 6 80

0.5

1

1.5

2

2.5

3

3.5

4

4.5

5x 10

4Thrust at Sea Level Static

Bypass Ratio

Th

rus

t (l

bf)

0 2 4 6 80

0.05

0.1

0.15

0.2

0.25

0.3

0.35

SFC at Sea Level Static

Bypass Ratio

SF

C (

lbm

/lb

f/h

r)

Prat=1.25

Prat=1.3

Prat=1.35

Prat=1.4

Prat=1.5

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

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

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

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

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

Weight BreakdownWing 5,200 lb

Fuselage 18,260 lb

Landing Gear 2,545 lb

Structure Total 27,000 lb

Engines 4,380 lb (x2)

Fuel 11,600 lb

Systems & Equipment 10,160 lb

Empty Weight 43,570 lb

Payload 2,730 lb

GTOW 58,700 lb

21

Wlanding gear, front

Wpassengers

WbaggageWcrew

Wfuel 2,3,4 A

Wfuel 1

Wlanding gear, main

Wverticle tail

Whorizontal tail

Wwing

Wengine Wfuselage

Weight LocationX 123 ft

61.6 ft

Wfuel 2,3,4 CWfuel 2,3,4 B

Neutral Point

22

Fuel Storage

12 ft

78 ft

35 ft

8 ftPax Area

3,42

1Passengers

243

1

1) D = 8 ft, L = 43 ft, V = 2027 ft3 8575 lb LH2

2) Each Section: D = 3 ft, L = 25 ft, V = 169 ft^3 717 lb LH2

Tank 2: V = 509 ft3 2152 lb LH2

3) Each Section: D = 1.5 ft, L = 25 ft, V = 41 ft3 172 lb LH2

Tank 3: V = 122 ft3 516 lb LH2

4) Each Section: D = 1.5 ft, L = 25 ft, V = 41 ft3 172 lb LH2

Tank 4: V = 122 ft3 516 lb LH2

Total: V = 2780 ft3 11760 lb LH2

Nose: 2*8 = 16 ft

Tail: 3.6*8 = 28 ft

Total: 78 + 16 + 24 = 123 ft

Fuel Weight = 11760 lb

LH2 Density = 4.23 lb/ft3

A B C

23

40000.00

42000.00

44000.00

46000.00

48000.00

50000.00

52000.00

54000.00

56000.00

58000.00

59.00 60.00 61.00 62.00 63.00 64.00 65.00

C.G. Location From Nose (ft)

We

igh

t (l

bs

)

Wo

Wo + Fuel

Wo + Fuel + All Cargo

End of Taxi & Take-off

End of Climb

End of Cruise

End of Decent

C.G. TravelNeutral Point

24

Dynamic Stability

Vertical Tail Volume

1 Engine Out Takeoff

350 ft2

Cross-wind Landing

307 ft2

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

HBJBBJ

A319

BJ

G550

Global

Exp

ress

Acquisition Cost ($2006 Millions) 60 52 45 38 39DOC ($2006 Dollars per hour) 4069 3200 2694 1820 1810

26

  HBJ BBJ G550GTOW (lbs) 58,700 171,000 91,000

Wing Area (ft2) 587 1345 1137

Span (ft) 83.53 117 91.5

Λ1/4c (Degree) 30 34 27

T/W 0.43 0.32 0.352

W/S (lbf/ft2) 100 127 80

Vcruise (KTAS) 460 450 460

Hcruise (ft) 40,000 39,000 40,000

Range (nmi) 5700 6200 6500

AR 11.9 10 7.4

(L/D)max 17 17.5 18.4

Landing Distance (ft) 3224 2549 2767

Takeoff Field Length (ft) 4459 5643 5934

Cost ($2006 Millions) 60 52 38

DOC ($2006) 4070 2900 1820

27

Outstanding Issues

• Stability

• Airfoil Selection and Aerodynamic Analysis

• Detailed Structural Analysis

• FAA Certification

• Research and Development Cost Analysis

28

Questions

29

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 Piercing Placed in Bonfire Charred Remains

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

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

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

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

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

LH2

36

Top View

8 ft

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

Constraint: Landing

39

Constraint: Landing

40

Constraint: Landing

41

Constraint: Landing

42

V-N Diagram Support

• Gust Velocities– At VB, G = 60.4 ft/s

– At VC, G = 45 ft/s

– At VD, G = 22.5 ft/s

• nGust = 1 + VG(KGGCLalpha)/(498W/S)

– KG = .88u/(5.3 + u)

– u = 2W/S/(p*cbar*g*CLalpha)

43

Fan Study at 40,000 ft

0 2 4 6 80

1000

2000

3000

4000

5000

6000

7000

8000

Thrust at Cruise Altitude

Bypass Ratio

Th

rus

t (l

bf)

0 2 4 6 80

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

SFC at Cruise Altitude

Bypass Ratio

SF

C (

lbm

/lb

f/h

r)

Prat=1.25

Prat=1.3

Prat=1.35

Prat=1.4

Prat=1.5

44

Engine Code Validation

Bypass Ratio

Pressure Ratio

Thrust (SLS)

Thrust (11 km)

SFC (11 km)

Cryoplane Engine 5.2 36.63 129.5 kN 24.77 kN

5.502 kg/(s-MN)

HBJ Engine Code 5.2 36.6 121.2 kN 33.3 kN

6.2 kg/(s-MN)

45

Landing Gear

• Size calculation– Used Business Twin equations– Table 11:1 – English Units

Diameter WidthA B A B

2.69 0.251 1.17 0.216

Diameter or Width (inches) = A * WwB

(where Ww is the weight applied on each wheel)

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

Spar Calculation

48

Stability Formulas

Add cg equ. Xbar=xi*Wi/Wi

Static margin formula = Xn-X bar/C bar

Equ for VHT

49

Component C.G. Location (ft)

Front Landing Gear 15

Crew 16

Baggage 25.5

Furnishings 33.5

Passengers 40

Miscellaneous 50

Fuel Tank 2,3,4 A 29.5

Fuel Tank 2,3,4 B 56.5

Fuel Tank 2,3,4 C 83.5

Fuselage 61.5

Component C.G. Location (ft)

Fuel Tank 1 74.5

Main Landing Gear 80

Wing 58.9

Nacelle 58.9

Engine 58.9

Vertical Tail 120

Horizontal Tail 121

C.G. Location

50

Stability Summary

 

C.G. (ft) SM Weight np-cg (ft)

Wo 61.25 16.84% 40827.00 1.30

Wo+1/2 Fuel 61.45 14.22% 42019.28 1.10

Wo+1/2 Fuel+crew 60.50 26.55% 42919.28 2.05

Wo+fuel 61.64 11.75% 43211.55 0.91

Wo+fuel+all cargo 61.74 10.57% 57687.97 0.82

Wo+all cargo 59.38 41.07% 43399.00 3.18

ready for take-off 61.74 10.57% 57687.97 0.82

taxi & take-off 61.73 10.67% 57627.00 0.83

climb 61.63 11.88% 56944.21 0.92

cruise 60.16 30.99% 47378.55 2.40

decent 60.12 31.40% 47260.55 2.43

51

Second Look at Stability

Wfuel 1

Wfuel 1Wfuel 5

Wfuel 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

Additional Trade Studies  Original 1/4 1/3

Wo 61.5 61.31 61.33

Wo+1/2 Fuel 62.47 61.37 61.09

Wo+1/2 Fuel+crew 62.03 60.63 60.35

Wo+fuel 63.95 61.62 61.10

Wo+fuel+all cargo 61.73 60.28 59.91

Wo+all cargo 62.42 57.93 60.05

ready for take-off 61.73 60.28 50.94

taxi & take-off 61.72 60.28 59.94

climb 61.62 60.26 59.95

cruise 60.13 59.99 60.07

decent 60.09 59.98 60.07

  Original 1/4 1/3

Wo 13.54 16.14 15.99

Wo+1/2 Fuel 1.03 15.45 19.00

Wo+1/2 Fuel+crew 6.67 25.11 28.68

Wo+fuel -18.16 12.17 18.89

Wo+fuel+all cargo 10.61 29.65 34.41

Wo+all cargo 1.64 27.63 32.59

ready for take-off 10.61 29.67 34.07

taxi & take-off 10.71 29.69 34.05

climb 12.07 29.95 33.95

cruise 31.33 33.39 32.31

decent 31.84 33.52 32.32

Center of Gravity (ft) Static Margin (%)

53

Drag Polar

54

1 Engine Out Calculations

• Fv = qv*Sv*CFβv*βv/β*δR

– qv (dynamic pressure at sea-level) = 48.61 slug/(ft*sec2)

– Sv (vertical tail area) ft2

– CFβv (tail lateral lift force coefficient) = 0.55

βv/β (free stream angle change) = 0.99

– δR (rudder deflection) = 0.35 radians

• T (thrust from 1 engine) = 12500 lbf

• M (total moment) = T*dE-Fv*dV = 0

– Calculate with distances from center of gravity. Solve for needed vertical tail area.

55

Cross-wind Landing

• Cnβ = Cnβw+Cnβfus+Cnβv-Fpβ/(q*Sw)* βv/β*(Xcg-Xp)

– Solve for Cnβv when Cnβ = 0

• Cnβv = CFβv*βv/β*ηv*Sv/Sw*(Xacv-Xcg)

– Solve for Sv to find required vertical tail area

56

Cost Support

• Historical Trend– $ = Awe

aMbRc

• 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

DOC Support

From Liebeck Cost Per Trip ($2006) Flight Crew 18798.83607

Airframe Labor 599.8911269Airframe Material Cost 441.9929672Maintenance Burden 2399.564508Engine Labor Cost 909.7189471

Engine Material Cost 669.0401125Engine Maintenance Burden 3638.875788

Landing Fee 64.64535Depreciation 17683.93009Insurance 1.39E+03

Fuel 7770Total Airframe Maintenance Cost 3441.448602Total Engine Maintenance Cost 5217.634848

Total per Departure 54365.99496Total Per Hour 4069.311

FLOPS DOC Per Departure CostTotal Maintenance and Labor 8884.61

Depreciation 17726.91Insurance 1839.68Servicing 27.28

Flight Crew 9092.53Fuel Cost 7768.63

Total 45339.64Total Per Hour 3393.685629

GTOW (lbs) 58,736Airframe Weight (lbs) 27,008

Thrust Per Engine (lbs) 12628# of Engines 2Fuel Weight 11,594Block Hours 13

Crew 4Period 14

Residual 0.01Airframe Spares 0.06Engine Spares 0.413

Airframe Cost ($) 3.97E+07# of Trips 200

58

DOC Support Contd

59

DOC Support Contd

60

DOC Support Cont