A I M 90-2317 Theoretical Performance of Unconventional Propellants and FuellOxidizer Combinations M. Rascon and K. Ramohalli Univ. of Arizona Tucson, AZ

AIAAISAEIASM EIASEE 26th Joint Propulsion Conference

July 16-18, 1990 1 Orlando, FL

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

THEORFIlCAL PERFORMANCE OF UNCONVENTIONAL PROPELLANTS AND FUEIJOXIDIZER COMBINATIONS

Mario Itascon* and Kumar ~amohalli** Aerospace and Mechanical Engineering University of Arizona, Tucson, AZ 85711

Abstract

Various fueUoxidizer and combinations are considered as propellants and their theoretical performance in rockets is compute over a wide range of pressures, expansion ratios and initial temperatures. The fundamental aim is to feed information for an intelligent choice to be made in space missions, where many of the more conventional (highly optimized) propellant combinations may not be the most advantages. For example, the much studied LOX/Hydrogen combination may deliver an excellent specific impulse but suffers from the necessity of refrigerated storage, access to hydrogen leaks. For some of the space missions that could utilize some of the local in-situ resources for manufacturing propellants, or of utilizing earth transported storables, the overall "figure of merit" appears to be superior to the more conventional cryogenic propellants. Continuing these thoughts, it turns out that several propellant combinations that may seem highly offdesign could be the best under specialized circumstances. Example include highly oxidizer-rich propellants on the moon.

In order to provide a complete data matrix to such future space missions, and possibly some terrestrial applications too, an extensive program was pursued. Various ingredients and combinations were run on four computer programs. The programs are: CET86 (NASA Lewis) mainframe version, a PC version of the CET86 (obtained from another University), a PC version of a program developed by a private company for specially considering charged species, and a PC version developed by another private company for specially handling large fractions of condensed (solid) species in the products. The outputs are seen to agree, normally within one percent in the chamber temperature. The

programs considered both equilibrium and frozen situations.

The propellants include conventional cryogenic and storables, and hydrogen peroxide, nitrogen tetroxide, LOX as oxidizers along with fuels that range from hydrogen to polymers (nylon, butyl rubber, etc.) that could be burned under special circumstances. The fueuoxidizer combinations cover a range of fuels and oxidizers that could be used in LEO, GEO, on Mars, the Moon, Phobos, Deimos and earth-crossing asteroids. Use of some "debris" is also cansidered.

These results were recently used in a paper dealing with Mars missions that employ in-situ resource utilization. That paper was given at the IAF congress in Spain.

Introduction

Space exploration/habitation is becoming important again. This time the USA will go to stay.12 It is obvious that the costs, with current technology, are prohibitive. Current launch vehicles consist mainly of propellants, with only a small fraction of the overall launch mass consisting of payload mass. This is illustrated in fig. 1.394

I U)I* -1-

Figure 1

** *NASA Traineeship Fellow, Student member, AIAA, Professor, Associate Fellow, AIAA

Copyright (c) 1990 Mario Rascon and Kumar Ramohalli. Published by thc American Institute of Aeronautics and Asuwautics, Inc. with permission.

Considering that various propellants and fuel/oxidizer combinations may become important for ISRU, it would be desirable to compute the theoretical performance of such combinations. The point is that the highly optimized conventional propellant combinations (LOX, Hydrogen for example) may not be available, or even the best choice for ISRU.~ Thus, three immediate regions of space were identified: near earth, Moon, Mars and Asteroids including Phobos and Deimos. Various propellant combinations are obvious in these regions. These are listed below. A newer thought in this arena is the possibility of using empty cases from previous stages as fuel; especially if the cases are filament wound with modern polymers. These filaments can be burned with ISRU oxidizers. Recall that the fuelloxidizer ratios are quite small.

The Choice of Propellants

For each of the three regions considered, propellants were determined by the availability of natural resources, and man carried resources, that could be used for the production of propellants. The aim is to find those propellant combinations that will give acceptable performance, such as ISp, but which present few, if any, storability problems and have the ability to use low technology hardware.

To this end propellants that may not give high performance values are considered, as well as unconventional chamber pressures, and expansion ratios. The choice of propellants is as follows:

Moon

Hz + 0 2 CH4 + 0 2 NH3 + N204 SiH4 + 0 2

H2 + Al + 0 2 Butyl Rubber + 0 2 Silicone + 0 2 Polystyrene + 0 2

For each of the propellants a wide range of O/F were considered in order to see what ratio would

be best to operate at. For some propellants it may best to operate at low O/F ratios since the oxidizer is the heaviest of the two. Operating at high O/F ratios could also be true if the fuel is the heaviest of the two.

The software used in this work is the Gordon and McBride CET86 program6 and the Air Force - Lockheed NOTS program.7 Two PC versions of the CET86 program were also obtained, one from the University of Minnesota, and the other from AVCO. These programs were used on a VAX- 8650 and an IBM PS 2 model 80 computers.

The CET86 program performs the calculations for performance parameters using standard thermodynamic equations. For actual computational calculations the program uses the Newton-Raphson iteration m e t h ~ d . ~ The assumptions used for calculation of performance parameters are: one-dimensional form of the continuity, momentum and energy equations; complete adiabatic combustion; zero velocity in the combustion chamber; isentropic expansion, homogeneous mixing; ideal gas law; zero temperature and velocity lags between condensed and gaseous species.10

All thermodynamic data is taken from a data file which include the JANAF thermodynamic set. This file is include with the CET86 program.

No source code or detailed documentation on the method of calculations used by the NOTS program was available since this program was obtained on a restricted use basis only. However, it is assumed that the method of calculations is not very different from that used by the CET86 program.

The programs were checked for proper behavior by comparing results to each other and to known values. In all cases the results were seen to agree, normally within one percent in the chamber temperature.

The software was tested for accuracy by comparing results to hand calculations. It was seen that the computer results compared quite nicely to those computed by hand. Hand calculations were done using the method described in ref. 11. As an example the CET86 was tested by using the reaction of methane and oxygen as a comparisons. The results were as follows:

Tmax (K) CO (moles) c o 2 H H2 H 2 0 NO

N2 0 O H 0 2

Hand 2233.0 0.1341 0.8659 ------ 0.0190 1.9474 ------ 7.5200 - - - - - - - 0.06921 0.05875

As can be seen the hand calculations and the CET86 results agree nicely within the limits of round-off error and the fact that H, 0 , and NO were not considered as possible product species in the hand calculations. The only reason these species were not considered was that it was felt that these species would have very small mole fractions and also would make the calculations easier.

Results

For a couple of propellants, plots of Isp vs. O/F at a constant pressure have been made to show the affect of chamber pressure on Isp. As can be seen, from figures 2 and 3, increasing the chamber pressure beyond 50 psia resulted in small gains

I

Figure 2: SiH,/LOX

\ P. = 1.0 psia

200 0 1 4

O/F

Figure 3: CO/LOX

Plots of Isp vs O F are also included for the propellants studied. The Isp(vac) values used are assuming equilibrium compositions. It is understood that the actual Isp values should be somewhere between equilibrium and frozen composition values. These graphs, figures 4 - 16, are simply to give a general idea of the ISp values obtainable for each propellant combmation. These graphs were made using the highest chamber pressure used with each of the propellant combinations along with four different expansion ratios considered.

Several propellants produce solid materials as a product of reaction which could cause problems, such as clotting and nozzle erosion. In all cases the solid materials were only significant at the lowest O/F ratios considered. Once a O/F ratio of about stoichiometric is reached the mole fractions of the solid species becomes insignificant. Plots of the mole fractions for the solid materials is included for these propellants in figures 17 - 20.

Further data on these propellan& is available in reference 12.

Discussion

All propellants combinations were processed to get their rocket performance for both equilibrium and frozen combinations during expansion. A large number of data was obtained on these propellant combinations, only a fraction of which is shown in this report. From the data it is seen that a number of these propellants give acceptable performance.

Of possible concern is the presence of solid material in the combustion products in certain propellant combinations. However, only the lowest Off ratios considered produced significant amounts.

Future Work

The purpose of this work was to provide a data base for not only the determination of possible propellants but also the determination of the best propellant for a particular mission. The reasoning is that one must not only consider the performance values of a particular propellant but also the availability, storability, etc. Also, one must consider what chamber pressure to use as well as the expansion ratio .to be used. Furthermore, each propellant may exhibit its own problems such as its combustion products and high temperature. All these variables, and more, must be balanced against each other in order to determine the optimum propellant and operating conditions for a particular mission. This could be accomplish by using weight factors that vary as their corresponding variables increase or decrease. For example, at some point increasing ISp further is not as important as keeping the chamber pressure and the expansion ratio down in order to minimize weight. The determination of these weight factors is presently being investigated at the University of Arizona.

References

Ride, Sally K: Leadership an America's Future in Space, A Report to the Administrator, August 1987.

Pain, Thomas 0.: The Next 40 Years in Space, IAF paper, 1989.

Jane's Spacecraft Directory, Jane's Publishing Inc., New York, NY 1987.

Logsdon, John M., and Williamson, Ray A: "U.S. Access to Space", Scientific American, v01. 260 NO. 3, p34-40.

Kirsh, Thomas A, and Ramohalli, Kumar N.R.: A "Figure of Merit" Approach to Extraterrestrial Resource Utilization. IAF paper 89-716. Oct. 1989.

Gordon, Sanford, and McBride, Bonnie J.: Computer Program for Calculation of

Complex Chemical EQuiliirium Cornpitions, Rocket Performance, Incident and Reflected Shocks, and Chapman-Jouguet Detonatins. Interim Revision. NASA SP-273, 1976.

AVCO Research Lab, 2385 Rever Beach Pkwy, Everett, MA 02149.

Naval Weapons Center, China Lake/Lockheed Propulsion Company Therrnochemical Rocket Motor Performance Prediction Program- NOTS, JPL.

Gordon, and McBride. hid.

Gordon, and McBride. p33-34. ibid

Strehlow, Roger A: Fundamentals of Combustion. Robert E. Krieger Publishing Co., p77-143.

Rascon, Mario H.: Unconventional Rocket Propellants. M.S. Report presented to the faculty of the University of Arizona, May 1990.

ISRU: In-Situ Resource Utilization.

OF: Oxidizer/Fuel ratio.

Specific Impulse.

P,: Chamber Pressure.

P,: Exit Pressure.

This work was sponsored by NASA Code XEU in support of the University of Arizona, Space Engineering Research Center. The authors thank Ms. Elaine Schwartz and John Lynch for financial support through the grant NGT. 70140.

""1 PC = 2000 psia

4001 PC = 2000 psia

'igure 5: NH,/N,04

PC = 100 psia

PC = 100 psia

400 - PC = 100 psi0

100 I , , , . , , , , , , , , , , , , , , , , , , , , , , , , , , , , 0 2 4 6 1

O F 8

igure 10 : Silicone/O,

400 1 P, = 100 psia

&

Figure 11 : Polystyrene/O,

PC = 500 psia

I

Figure 12: CO/O,

400 PC = 500 psia

zoo

igure 13: CO/H,O,

PJP = 1 WO 0 PJP: = 100

P / P = l o 0 P:/P: = 2

I I Figure 15 : C,H,OH/O,

1 1 PC = 500 psia

Figure 17: SiHJO,

3.50 + CARBON CONCENTRATION 0/F = 1.34

3 1 W a : 0 3 0 0 - (L a

ri I c 3 "'2.50 - 0 - 0 - \ 1

cn 2.00 - w - 1 : 0 - I

1.50

PRESSURE(psio)

CARBON CONCENTRATION O/F = 1.36

a 0 CY a

- 0.00

PRESSURE(~S~O)

Figure 19 : ~ilicone/O,

CARBON CONCENTRATION O/F = 2.15

0.00 1:o

P R E S S U R E ( ~ S ~ ~ ) L

Figure 2 0 : Polystyrene/O,

E C f b QLE

LEO 100 - 2000 2 - 1000 .4 - 170

MOON .1 - 100 2 - lo00 .6 - 30

MARS 1 - 500 2 - lo00 .3 - 34

Total of over 50000 data points

Table 1

Figure 18: B-R/02