Thermal Engines for Launch Vehicle Configurations

February 18, 2006 HYPERIONERAU

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Thermal Engines for Launch Vehicle Configurations


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Agenda

• What is propulsion

• Thermal engine basics

• LOX Augmented Thermal Engines

• Launch Vehicle Dynamics

• SSTO/MSTO


3

Propulsion

Propulsion is energy

Energy and momentum arerelated

more energy = more propulsion


4

Energy Sources

Fuels Energy Release J/kg Converted Mass Fraction

Chemical

LO/LH 1.35 x 107 1.25 x 10-10

Atomic Hydrogen 2.18 x 108 2.40 x 10-9

Metastable Helium 4.77 x 108 5.30 x 10-9

Nuclear Fission238U 8.20 x 1013 9.10 x 10-4

Nuclear Fusion

DT (0.4/0.6) 3.38 x 1014 3.75 x 10-3

CAT-DT (1.0) 3.45 x 1014 3.84 x 10-3

D3He (0.4/0.6) 3.52 x 1014 8.90 x 10-3

pB11 (0.1/0.9) 7.32 x 1013 8.10 x 10-4

Matter-Antimatter 9.00 x 1016 1


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Thermal Engines

• Produce heat to expand a propellant

• Requires core materials to withstand high melting points (~2000-4000 K) such as Tungsten and Carbon

• Propellant must have a high Cp and low atomic weight. Liquid Hydrogen is a primary candidate


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Specific Impulse

Specific Impulse Vs. Core Temperature

0

200

400

600

800

1000

1200

1400

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000

Core Temperature (K)

Spec

ific

Impuls

e (s

)

Series1


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Thrust

Thrust per Unit Area of Transfer Surface Vs. Core Temperature

0

20

40

60

80

100

120

140

160

180

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000


Thru

st p

er U

nit A

rea

of C

onve

ctiv

e Tra

nsf

er S

urf

ace

(N/m̂

2)

Series1


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Nuclear Thermal Rocketry (NTR)

-Heavily tested-Must use a combination of W, LiH, Be-Radiation and spallation-Limited by material melting point


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Positron Thermal Rocket (PTR)

• Identical to NTR

• Uses positrons as a heat source

• Concentric cylinder configuration

• Requires only Tungsten, allowing higher core temp. and Isp


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LOX Augmentation

Increases thrust

Decreases Isp


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LOX Augmentation

T (K) T (K) T (K) T/W

O/F 2900 2800 2600

0 941 925 891 3

1 772 762 741 4.8

3 647 642 631 8.2

5 576 573 566 11

7 514 512 508 13.1

- Based on NTR NERVA configuration-Assumes that PTR is scaled to NTR NERVA specs.- More advanced PBR could increase T/W by X7


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LOX Augmentation

Specific Impulse Vs. O/F Ratio

0

200

400

600

800

1000

1200

0 1 2 3 4 5 6 7 8

O/F Ratio

Sp

ecif

ic Im

pu

lse

(s)

2900 K

2800 K

2600 K

3300 K


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LOX Augmentation

% Decrease in Isp Vs. O/F Ratio

0

10

20

30

40

50

60

0 1 2 3 4 5 6 7 8

O/F Ratio

% D

ecre

as

e in

Is

p

2900 K

2800 K

2600 K

3300 K


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LOX Augmentation

% Decrease in Specific Impulse Vs. Temperature

y = 0.0038x + 7.0279

y = 0.0069x + 11.241

y = 0.0077x + 16.373

y = 0.008x + 22.143

0

5

10

15

20

25

30

35

40

45

50

2550 2600 2650 2700 2750 2800 2850 2900 2950


% D

ecre

ase

in Is

p

O/F = 1

O/F = 3

O/F = 5

O/F = 7

Linear (O/F = 1)

Linear (O/F = 3)

Linear (O/F = 5)

Linear (O/F = 7)


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LOX Augmentation

T (K) T/W

O/F 3300

0 988 3

1 794.6691 4.8

3 651.9713 8.2

5 575.184 11

7 508.3952 13.1


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Launch Vehicle

f

isp M

MgIV ln

rVo


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Launch Vehicle Requirements

rotationogravitydragsteeringtotal VVVVVV

rotationob

i

i

i

bitotal VVgt

M

tmM

m

D

M

tmM

m

FV

sinlnlncos1

sin2

lnsin

lncos1sin

lnsin

2

2

22

tgtm

M

tmMtmM

m

D

tmM

tmMtmM

m

Ftm

M

tmMtmM

m

FH

i

ii

i

ii

i

ii


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Single Stage to Orbit (SSTO)

• Simple

• Quick turnaround time

• Short loiter time in orbit

• Small payloads delivered


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SSTO

Specific Impulse Vs. Payload Mass Fraction to LEO for a SSTO Rocket (assume ΔV = 10.00 km/s)

0

200

400

600

800

1000

1200

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45

Payload Mass Fraction

Spec

ific

Impuls

e (s

)

Series1


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SSTO

Payload Mass Fraction to LEO Vs. O/F Ratio (assumes reactor temperature of 3300 K and a ΔV of 9.00 km/s)

0

50

100

150

200

250

300

LOX/LH2 7 5 3 1 0

O/F Ratio

Mas

s Fra

ctio

n /

% In

crea

se

Payload Mass Fraction

% Increase


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Multistage Rocketry (MSTO)

• Assume all 1st stage engines are Saturn F-1’s

Isp: 330 seconds

T/W: 96

• Upper-stage chemical engines are SSME’s

Isp; 450 seconds

T/W: 73


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MSTO Assumptions

1 2 3

Drag (kN) 100 50 0.2

alpha (degrees) 5 1 0

gamma (degrees) 60 30 10


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MSTO

Payload to LEO vs. Vehicle Configuration (assumes 300 km circular orbit)

0

20

40

60

80

100

120

140

SSME/SSME 652TE/652TE 652TE/988TE 652TE/794TE SSME/784TE SSME/794TE SSME/988TE

Vehicle Configuration

Pal

yoad

to

Orb

it (

Mt)

Payload to LEO


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MSTO

Mass Fraction to LEO vs. Vehicle Configuration (assumes a 300 km circular orbit)

0

2

4

6

8

10

12



Pay

load

Mas

s F

ract

ion

Mass Fraction to LEO


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MSTO

% increase in Payload Mass Fraction to LEO (assuming a 300 km circular orbit)

0

5

10

15

20

25

30



% In

crea

se

% increase


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MSTO


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MSTO


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Future Work

• Validate Altitude equation

• Consider using a PBR analysis

• Use Mars transfer analysis for baseline PTR configuration

• Use more precise computer model to demonstrate changes in D, alpha, gamma, etc…

• Determine launch cost to include H2, O2, etc…


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References

1. Blevins, J., Patton, B., Ryhs, N., Schmidt, G., Limitations of Nuclear Propulsion for Earth to Orbit, AIAA Paper 2001-3515

2. Smith, D., Wulff, J., Pearce, C., Bingaman, J., Webb, J., Thermal Radiation Studies for an Electron Positron Annihilation Propulsion System, AIAA Paper 2005-3230

3. Humble, R., Henry, G., Wiley, J., Space Propulsion Analysis and Design, McGraw Hills Co. Inc. 1995

4. Smith, G., Kramer, K., Meyer, K., Thode, T., High Density Storage of Antimatter for Space Propulsion Application, AIAA Paper 2001-3230

5. Borowski, S., Dudzinski, L., 2001 A Space Odyssey Revisited – The Feasibility of 24 Hour Commuter Flights to the Moon Using NTR Propulsion with LUNOX Afterburners, published with permission from NASA, AIAA Paper 97-2956

6. Bulman, M., Messit, D., Niel, T., Borowski, S., High Area Ratio LOX-Augmented Nuclear Thermal Rocket (LANTR) Testing, AIAA Paper 2001-3369


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Questions/Comments

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Thermal Engines for Launch Vehicle Configurations