Spacecraft Propellants

(Updated 17 February 2012)

References

Rocket Propulsion Analysis (Lite Edition) v.1.2.5 by www.propulsion-analysis.com

Note on ISP

You might wonder how the RL-10B-2 which has an ISP of 464 seconds, can have a higher ISP than the commonly quoted theoretical maximum of 451-455 seconds for LOX/LH2 found on the internet.

Or how documents which list theoretical ISPs for different types of fuels can have wildly differing numbers, e.g. Document A says LOX/LH2 has a vacuum ISP of 454, while Document B says it’s 462.

The answer is that the maximum theoretical ISP of a propellant is heavily dependent on engine-specific numbers like chamber pressure (PC), nozzle expansion ratio (Ɛ), and the Oxidizer/Fuel Ratio (O/F). If you use different numbers, you get different theoretical vacuum ISPs.

In the example above, Document A had a PC of 1000 psia, Ɛ=40 and O/F=4.83, while Document B had a PC of 1000 psia, Ɛ=60 and O/F=5.0

Expansion ratio (Ɛ) in rocketry is the area of the nozzle exit divided by the area of the nozzle throat. If your engine does all of it’s work in an atmosphere, you want a low expansion ratio to provide good sea-level ISP (the F-1’s Ɛ=16 and the RS-68’s Ɛ=21.5), while if your engine lives in vacuum, you want a large expansion ratio (the RL-10A-1’s Ɛ=40 and the J-2X’s Ɛ=90).

At the far edge of currently proved nozzle technology is the RL-10B-2’s astounding Ɛ=285, which is achieved through a carbon-carbon nozzle extension which is stowed in a “retracted” position to take up less space during launch (Cutaway photo of RL-10B-2 in stowed mode). After stage separation, the Nozzle Extension Deployment System (NEDS) uses a series of ball-screw drives to lower the nozzle extension over the core engine nozzle.

Chamber Pressure (PC) has very little effect on theoretical vacuum ISP, but a large effect on sea level ISP. To illustrate this point; we’ll use the Space Shuttle Main Engine (SSME). It has Ɛ=77.5; which with an O/F ratio of 6.03, gives us the following theoretical maximum specific impulse:

PC: 1000 psia = Sea Level 196.52 seconds, Vacuum 463.77 seconds
PC: 3000 psia = Sea Level 374.88 seconds, Vacuum 464.69 seconds

You can see how a very high PC allows for a significant improvement in sea level performance for nozzles with a large Ɛ; albeit at increased cost and complexity. Currently, the highest production PC is found on Russian staged combustion kerosene engines such as the RD-180 and RD-191, which reach 3,500 to 3,800 psia.

Oxidizer/Fuel Ratio (O/F): By varying the ratio of oxidant, you can control ISP and chamber temperature as needed. A good example would be this graph which was calculated through Rocket Propulsion Analysis (Lite Edition) v.1.2.5.

As your expansion ratio changes, so does the O/F mixture needed to achieve optimum performance, as shown by this graph which was also calculated through RPA (Lite Edition) v1.2.5.

It’s worth noting that these graphs were calculated with the assumption of a perfect combustion process; e.g. every molecule of fuel is perfectly combusted with it’s counterpart oxidant molecule, which doesn’t happen in reality. This is most notable with liquid hydrogen fuelled engines.

Calculating Overall Propellant Density

PropDens = (OFR + 1) / [ (OFR / OxidizerDensity) + (1/FuelDensity) ]

Where:

OFR: Oxidizer/Fuel Ratio. If it is 6 [to 1], then put in 6.
OxidizerDensity: Density of oxidizer in kg/m3
FuelDensity: Density of fuel in kg/m3
PropDensity: Density of propellant in kg/m3

Calculating Metallized Gel Propellant Density

PropDens = 1 / [ (1 – ML – GL) / LPD + (AL / AD) + (GL / GD) ]

Where:

LPD: Liquid Propellant Density
AD: Additive Density
AL:
Additive Loading
GD: Gellant Density
GL: Gellant Loading

EXAMPLE: The density of a metallized gel propellant that used liquid hydrogen as a base (71 kg/m3) with 60% by weight Aluminum additive (2,768 kg/m3) and 10% by weight of Methane (CH4) gellant (520 kg/m3) would be:

1 / [ (1 – 0.6 – 0.1) / 71 + (0.6 / 2768) + (0.1 / 520) ] = 213 kg/m3

Standard Propellant Combinations
(Widely understood, and easily handleable)

NOTE: The optimum O/F ratio depends on the surrounding environment – the optimum propellant mix for a engine is different between vacuum and sea level – and generates different levels of ISP. The mixtures below were formulated for optimum vacuum ISP, at the loss of some sea level ISP.

Hydrogen (LH2) / Liquid Oxygen (LOX)

Actual Reference O/F Ratios: 6.0 (SSME), 5.85 (RL-10B-2), 5.50 (J-2X).
Actual Engine Efficiencies: 97.7% (SSME), 96.6% (RL-10B-2), 95.7% (J-2X)

Theoretical Optimum Performance

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

Mixture Ratio
(O/F)

Prop.
Density (kg/m3)

ISP
(s/l)

ISP
(vac)

1000

20 to 1
(RS-68)

4.323

297.76

369.4

441.48

40 to 1
(Std. Ɛ)

4.645

310.86

311.78

455.06

90 to 1
(J-2X)

4.965

323.49

147.52

467.77

285 to 1
(RL-10B-2)

5.517

344.38

N/A

481.66

3000

20 to 1
(RS-68)

4.396

300.77

417.74

441.79

40 to 1
(Standard Ɛ)

4.726

314.1

407.59

455.42

90 to 1
(J-2X)

5.046

326.62

361.15

468.18

285 to 1
(RL-10B-2)

5.563

346.07

147.31

482.1

Generalized Actual Performance
O/F = 6.0, density of 361.8 kg/m3

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

ISP
(s/l)

ISP
(vac)

1000

20 to 1
(RS-68)

366.74

435.78

40 to 1
(Std. Ɛ)

313.60

451.66

90 to 1
(J-2X)

155.66

466.31

285 to 1
(RL-10B-2)

N/A

481.58

3000

20 to 1
(RS-68)

413.96

437.16

40 to 1
(Standard Ɛ)

406.38

452.77

90 to 1
(J-2X)

362.78

467.16

285 to 1
(RL-10B-2)

151.37

482.17

Kerosene (RP-1) / Liquid Oxygen (LOX)

Actual Reference O/F Ratios: 2.72 (RD-180), 2.27 (F-1), 2.25 (MA-5A), 2.23 (H-1)
Actual Engine Efficiencies: 94.2% (RD-180), 92.1% (MA-5A) 91.8% (H-1), 90.1% (F-1)

Theoretical Optimum Performance

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

Mixture Ratio
(O/F)

Prop.
Density (kg/m3)

ISP
(s/l)

ISP
(vac)

1000

20 to 1
(RS-68)

2.635

1023.34

289.72

343.37

40 to 1
(Std. Ɛ)

2.715

1025.60

250.73

357.64

90 to 1
(J-2X)

2.812

1028.22

132.46

371.88

285 to 1
(RL-10B-2)

2.932

1031.31

N/A

388.62

3000

20 to 1
(RS-68)

2.699

1025.15

327.73

345.78

40 to 1
(Standard Ɛ)

2.776

1027.26

324.04

360.02

90 to 1
(J-2X)

2.863

1029.56

293.55

374.18

285 to 1
(RL-10B-2)

2.974

1032.35

136.89

390.77

Generalized Actual Performance
O/F = 2.72, density of 1025.73 kg/m3

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

ISP
(s/l)

ISP
(vac)

1000

20 to 1
(RS-68)

289.82

343.25

40 to 1
(Std. Ɛ)

250.76

357.65

90 to 1
(J-2X)

130.84

371.34

285 to 1
(RL-10B-2)

N/A

386.37

3000

20 to 1
(RS-68)

327.77

345.8

40 to 1
(Standard Ɛ)

323.71

359.78

90 to 1
(J-2X)

291.9

373.05

285 to 1
(RL-10B-2)

130.64

387.64

Monomethyl Hydrazine (MMH) / Nitrogen Tetroxide ( NTO)

Notes: This is currently the “standard” storable hypergolic propellant, though attempts are underway to replace it with ‘green’ non-toxic propellants.

Due to it being used heavily for space storable systems, and also because the Soviet Union designed a 3,140~ psia PC storable propellant engine, the RD-264 for the R-36 (SS-18) ICBM, the table has been expanded to add 150 PC.

The 150 psia PC compares quite well to the 100 psia found on the Apollo Service Propulsion System, 116~ psia on the Lunar Module Descent engine, and the 125 psia on the Space Shuttle’s Orbital Maneuvering System.

Actual Reference O/F Ratios: 1.65 (Shuttle OMS)
Actual Engine Efficiencies: 93.4% (Shuttle OMS)

Theoretical Optimum Performance

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

Mixture Ratio
(O/F)

Prop.
Density (kg/m3)

ISP
(s/l)

ISP
(vac)

150

20 to 1
(RS-68)

2.067

1188.64

N/A

326.48

40 to 1
(Std. Ɛ)

2.134

1193.24

N/A

338.66

90 to 1
(J-2X)

2.181

1196.37

N/A

350.19

285 to 1
(RL-10B-2)

2.446

1212.68

N/A

364.42

1000

20 to 1
(RS-68)

2.126

1192.70

276.60

328.94

40 to 1
(Standard Ɛ)

2.180

1196.31

236.5

340.92

90 to 1
(J-2X)

2.185

1196.63

116.8

351.83

285 to 1
(RL-10B-2)

2.473

1214.23

N/A

366.23

3000

20 to 1
(RS-68)

2.174

1195.91

312.67

330.23

40 to 1
(Standard Ɛ)

2.181

1196.37

306.68

341.80

90 to 1
(J-2X)

2.423

1211.35

276.49

354.39

285 to 1
(RL-10B-2)

2.480

1214.63

121.19

367.12

Generalized Actual Performance
O/F = 1.65, density of 1155.86 kg/m3

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

ISP
(s/l)

ISP
(vac)

150

20 to 1
(RS-68)

N/A

321.11

40 to 1
(Std. Ɛ)

N/A

331.20

90 to 1
(J-2X)

N/A

340.51

285 to 1
(RL-10B-2)

N/A

350.34

1000

20 to 1
(RS-68)

269.24

321.84

40 to 1
(Standard Ɛ)

226.57

331.78

90 to 1
(J-2X)

104.23

340.95

285 to 1
(RL-10B-2)

N/A

350.83

3000

20 to 1
(RS-68)

304.51

322.09

40 to 1
(Standard Ɛ)

296.82

331.98

90 to 1
(J-2X)

262.01

341.11

285 to 1
(RL-10B-2)

100.65

351.15

Exotic Propellant Combinations
(Not understood well, and/or extremely dangerous to handle)

Notes: Due to the largely experimental nature of these propellants, you are encouraged to be strongly conservative in picking these for your hypothetical spacecraft via using conservative engine efficiency ratios (around 94% or so).

Aluminum/Hydrogen Cryogel / Liquid Oxygen (LOX)

Notes: The cryogel’s density is 170.9 kg/m3; of which the aluminum additive is 60% by weight.

While this system actually is slightly less dense than straight LH2/LOX, there’s a massive beneficial spiral of weight decreases due to the massively increased density of the fuel (71 to 170.9 kg/m3) and the lower oxidizer to fuel ratio.

A crude estimate of a S-II sized system holding 72,780 kg of fuel has a total system mass of 525.13 tonnes for the ‘straight’ LH2 version, and 195.24 tonnes for the Aluminum/Hydrogen Cryogel version, while delta-v remains largely the same.

Reference: NASA TM-1998-206306: Preliminary Assessment of Using Gelled and Hybrid Propellant Propulsion for VTOL/SSTO Launch Systems.

Theoretical Performance
O/F = 1.6, density of 358.4 kg/m3

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

ISP
(s/l)

ISP
(vac)

1000

20 to 1
(RS-68)

370.07

439.62

40 to 1
(Standard Ɛ)

316.92

456.04

90 to 1
(J-2X)

158.26

471.25

285 to 1
(RL-10B-2)

N/A

487.3

3000

20 to 1
(RS-68)

417.14

440.44

40 to 1
(Standard Ɛ)

410.11

456.7

90 to 1
(J-2X)

366.92

471.76

285 to 1
(RL-10B-2)

155.39

487.66

Hydrogen (LH2) / Fluorine (LF2)

Notes: You might wonder why Liquid Fluorine isn’t used more. The answer is that Fluorine is violently unstable and likes to spontaneously combust on contact with such sundry things as the air, or rocket engineers.

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

Mixture Ratio
(O/F)

Prop.
Density (kg/m3)

ISP
(s/l)

ISP
(vac)

1000

20 to 1
(RS-68)

9.272

507.89

390.41

466.41

40 to 1
(Std. Ɛ)

10.721

553.28

329.06

480.01

90 to 1
(J-2X)

12.250

596.93

154.90

492.54

285 to 1
(RL-10B-2)

13.722

635.34

N/A

505.31

3000

20 to 1
(RS-68)

10.007

531.44

442.75

468.26

40 to 1
(Standard Ɛ)

11.465

575.03

431.12

481.85

90 to 1
(J-2X)

13.138

620.49

380.81

494.30

285 to 1
(RL-10B-2)

14.364

651.10

149.27

506.70

Monomethyl Hydrazine (MMH) / Fluorine (LF2)

Notes: This mixture was the “last man standing” out of JPL/NASA/USAF studies conducted from the 1960s to the late 1970s for a next generation space-storable propellant. The losing mixtures were Oxygen Difluoride (OF2) / Diborane (B2H6), which was studied from 1966-1971 and various fluorine-oxygen mixtures (FLOX) combined with MMH.

Due to this essentially being designed for a space storable system, and also because the Soviet Union designed a 3,140~ psia PC storable propellant engine, the RD-264 for the R-36 (SS-18) ICBM, the table has been expanded to add 150 PC.

The 150 psia PC compares quite well to the 100 psia found on the Apollo Service Propulsion System, 116~ psia on the Lunar Module Descent engine, and the 125 psia on the Space Shuttle’s Orbital Maneuvering System.

Typical efficiency of this kind of mixture in a practical engine would be about 93%.

References:
Experience with Fluorine and its Safe Use as a Propellant – NASA JPL Publication 79-64 (30 June 1979)

Chamber Pressure
(Pc)

Aspect
Ratio
(Ɛ)

Mixture Ratio
(O/F)

Prop.
Density (kg/m3)

ISP
(s/l)

ISP
(vac)

150

20 to 1
(RS-68)

2.427

1240.76

N/A

392.92

40 to 1
(Std. Ɛ)

2.459

1242.81

N/A

410.07

90 to 1
(J-2X)

2.472

1243.63

N/A

425.94

285 to 1
(RL-10B-2)

2.473

1243.7

N/A

442.06

1000

20 to 1
(RS-68)

2.456

1242.62

336.56

398.49

40 to 1
(Standard Ɛ)

2.470

1243.51

291.10

414.82

90 to 1
(J-2X)

2.473

1243.70

151.18

429.61

285 to 1
(RL-10B-2)

2.478

1244.01

N/A

444.47

3000

20 to 1
(RS-68)

2.465

1243.19

380.01

400.85

40 to 1
(Standard Ɛ)

2.472

1243.63

375.11

416.76

90 to 1
(J-2X)

2.473

1243.70

337.36

431.07

285 to 1
(RL-10B-2)

2.511

1246.07

148.46

445.19

Oxidizers (By Density)

Fluorine (LF2)

Density (Liquid): 1,510 kg/m3
Freezing Point: -219.6 C
Boiling Point: -188.1 C

Oxygen (LOX)

Density (Liquid): 1,140 kg/m3
Freezing Point: -219 C
Boiling Point: -183 C

Nitrogen Tetroxide / NTO (N2O4)

Density (Liquid): 1450 kg/m3
Freezing Point: -9.3 C
Boiling Point: 21.15 C

Fuels (By Density)

Monomethyl Hydrazine / MMH (CH3NHNH2)

Density (Liquid): 866 kg/m3
Freezing Point: -52.4 C
Boiling Point: 87.5 C

RP-1 Kerosene

Density (Liquid): 806 kg/m3
Freezing Point: -73 C
Boiling Point: 147 C

Hydrogen (LH2)

Density (Liquid): 71 kg/m3
Freezing Point: -259 C
Boiling Point: -253 C

Solid Propellants

TP-H1148

Type: PBAN
Burn Rate: 0.42 to 0.47 inches per second at 1000 psi
Density: 1,757.67 kg/m3
Iosps: 261.9 using HPM nozzle
ISP (vac): 268.5 using HPM nozzle
Composition:
      PBAN/Epoxy Binder: 14%
      Iron Oxide (Fe2O3): 0.28% (varied throughout mix to control burn rate)
      Aluminum Powder: 16%
      Aluminum Perchlorate: 69.72% (varied throughout mix to control burn rate)
      Total Solids: 86%
Toxic Exhaust Products
      HCl: 0.5860 moles per 100 grams
      Al2O3: 0.2965(S) moles per 100 grams
      CO: 0.8555 moles per 100 grams
Notes: Used in Space Shuttle High Performance Motor. Meets the Space Shuttle HPM Specification, which requires it to maintain integrity over a five year storage period between 35-95F.
References:
Block II Solid Rocket Motor (SRM) Conceptual Design Study (NAS 8-37295) Final Report Vol 1: Appendices by Atlantic Research Corporation (31 Dec 1986) (PDF excerpt)
Block II SRM Conceptual Design Studies Final Report: Conceptual Design Package: Volume I, Book 1 – by Morton Thiokol (19 Dec 1986) (PDF Excerpt)

DL-H396 Propellant

Type: HTPB
Density: 1,801.4 kg/m3
ISP (vac): 280.04 at 7.72 Expansion Ratio and 1000 psia
Composition (by percentage of weight):
      R-45HT/HTPB Polymer: 11.02%
      HX-752 Aziridine Bonding Agent: 0.15%
      Iron Oxide (Fe2O3): 0.2%
      Aluminum Powder: 19%
      Aluminum Perchlorate (200/20 microns): 68.92%
      IDPI Curing Agent: 0.71%
      Total Solids: 88%
Notes: Modified Peacekeeper Stage I formulation using lower cost ingredients such as non-spherical aluminum and RT-45HT/HTPB polymer, which costs about 50% less than PBAN or R-45M/HTPB polymer.
References:
Block II SRM Conceptual Design Studies Final Report – Conceptual Design Package: Volume I, Book 1 – Morton Thiokol (19 December 1986) (PDF excerpt with specifications)
Block II SRM Conceptual Design Studies Final Report – Preliminary Development and Verification Plan: Volume I, Book 2 – Morton Thiokol (19 December 1986) (PDF excerpt)

DL-H397 Propellant

Type: HTPB
Composition (by percentage of weight):
      HTPB/Isocyanate Binder: 10%
      DOA: 2%
      Iron Oxide (Fe2O3): 0.25% (varied for burn rate control)
      Aluminum Powder: 19%
      Aluminum Perchlorate: 39.50%
      Sodium Nitrate (NaNO3): 0.25% (varied for burn rate control)
      Aluminum Oxide (Al2O3): 0.25% (varied for burn rate control)
Notes: Designed for very low HCl exhaust products (less than 1%).
References:
Block II SRM Conceptual Design Studies Final Report – Conceptual Design Package: Volume I, Book 1 – Morton Thiokol (19 December 1986) (PDF excerpt with specifications)
Block II SRM Conceptual Design Studies Final Report – Preliminary Development and Verification Plan: Volume I, Book 2 – Morton Thiokol (19 December 1986) (PDF excerpt)

TP-H3340 Propellant

Type: HTPB
Density: 1,815.8 kg/m3
ISP: 280.11 (unknown what is used as baseline to establish this)
Composition:
      HTPB/Isocyanate Binder: 11% (HX-752 Aziridine bonding agent)
      Aluminum Powder: 18%
      Aluminum Perchlorate: 71%
      Total Solids: 89%
Notes: Used in Star 37X Motor.
References:
Block II SRM Conceptual Design Studies Final Report – Conceptual Design Package: Volume I, Book 1 – Morton Thiokol (19 December 1986) (PDF excerpt with specifications)
Block II SRM Conceptual Design Studies Final Report – Preliminary Development and Verification Plan: Volume I, Book 2 – Morton Thiokol (19 December 1986) (PDF excerpt)

ARCADENE 360A

Type: HTPB
Density: 1,813.03 kg/m3
Iosps: 260.7 lbf-sec/lbm
Composition (by percentage of weight):
      R-45 HT Binder: 10%
      DOA: 2%
      Aluminum Powder: 18%
      Iron Oxide (Fe2O3): 1.5
      Aluminum Perchlorate (60/40 200 μ/MA): 68.5%
      Total Solids: 88%
Notes: Variant of 1980s MLRS propellant proposed for use as SRB Igniter.
References:
Block II Solid Rocket Motor (SRM) Conceptual Design Study (NAS 8-37295) Final Report Vol 1: Appendices by Atlantic Research Corporation (31 Dec 1986) (PDF excerpt)

ARCADENE 360B

Type: HTPB
Binder: R-45HT/IPDI
Density: 1,799.19 kg/m3
Iosps: 263.1
ISP (vac): 269.7
Composition:
      Iron Oxide (Fe2O3): 0.1% to 0.3%
      Aluminum Powder: 18%
      Total Solids: 88%
      Aluminum Perchlorate Mix: 70% Coarse, 30% Fine
Toxic Exhaust Products
      HCl: 0.5837 moles per 100 grams
      Al2O3: 0.2989(S) moles per 100 grams
      CO: 0.7697 moles per 100 grams
Notes: Was being produced at rates of 70,000 lb/day for the Vought MLRS program around 1986. ISP is designed around using the Space Shuttle High Performance Motor as a baseline.
References:
Block II Solid Rocket Motor (SRM) Conceptual Design Study (NAS 8-37295) Final Report Vol 1: Appendices by Atlantic Research Corporation (31 Dec 1986) (PDF excerpt)