/I/7 i .:i_// it," .C_._ _:.....C:_J EVALUATION OF ON-BOARD HYDROGEN STORAGE METHODS FOR HIGH-SPEED AIRCRAFT (NAG-I-7 67) STATUS REPORT January 31, 1991 INVESTIGATORS Ates Akyurtlu Jale F. Akyurtlu Hampton University Department of Engineering _;_?1.- 2 2....... ...... "-_!_77!_,5) :=4A_ .., <:_ .1__:,_._,i_. /_,_ _ .: _.:_ L_J_ ,......... " " "......... Ui']C ] :.:__ G310 5 O0 12 3__% TABLE OF CONTENTS ABSTRACT .............. 1 INTRODUCTION ............. 2 EVALUATION CRITERIA .......... 2 MODELING CONSIDERATIONS ........ 4 RESULTS ............. 7 RECOMMENDATIONS ........... ii REFERENCES .............. 13 APPENDIX A APPENDIX B APPENDIX C APPENDIX D APPENDIX E ABSTRACT Hydrogen is the fuel of choice for hypersonic vehicles. Its main disadvantage is its low liquid and solid density. This increases the vehicle volume and hence the drag losses during atmospheric flight. In addition, the dry mass of the vehicle is larger due to larger vehicle structure and fuel tankage. Therefore it is very desirable to find a fuel system with smaller fuel storage requirements without deteriorating the vehicle performance substantially. To evaluate various candidate fuel systems, they were first screened thermodynamically with respect to their energy content and cooling capacities. To evaluate the vehicle performance with different fuel systems, a simple computer model is developed to compute the vehicle parameters such as the vehicle volume, dry mass, effective specific impulse, and payload capacity. The results indicate that if the payload capacity (or the gross lift-off mass) is the most important criterion, only slush hydrogen and liquid hydrogen - liquid methane gel shows better performance than the liquid hydrogen vehicle. But if all the advantages of a smaller vehicle is considered and a more accurate mass analysis can be performed, other systems using endothermic fuels such as cyclohexane, and some boranes may prove to be worthy of further consideration. 1 INTRODUCTION Hydrogen is a foremost candidate as a fuel for use in hypersonic flight. The National Aerospace program has been initiated by NASA ad the Department of Defense for developing hypersonic / transat- mospheric vehicles for takeoff from conventional airport runways to orbit, or for rapid, long distance, intercontinental aerospace transportation. For this purpose, air-breathing, hydrogen fueled supersonic combustion ramjet (scramjet) engines are being developed for speeds of Mach 5 to 25. The main difficulty encountered in the use of hydrogen as a fuel for hypersonic vehicles is the large volume required for its on-board storage. If hydrogen is stored as liquid, it requires about four times the volume to produce the same amount of energy as conventional fuels. This is especially important for supersonic and hypersonic vehicles which need to have slender designs to reduce drag losses. Initially the objective of this study was to identify and evaluate the storage media capable of increasing the hydrogen storage density (mass of hydrogen stored per unit storage volume) to a level _ higher than that of liquid hydrogen (70 kg/m3). It was then realized that since the fuel system and the vehicle formed a complex system, any improvement in the hydrogen density would involve several trade-offs. For this reason, the establishment of a set of criteria for the evaluation of various fuel systems and putting together a model which will make a quantitative evaluation possible became the primary objective of this work. EVALUATION CRITERIA During hypersonic flight, beside providing propulsion, the fuel has to contribute to structural and engine cooling. In addition, combustibles other than hydrogen in the fuel system may serve as rocket fuel during the final stage of flight to orbit and for maneuvering in space, or they may be burnt to provide power for the vehicle subsystems. Therefore, the hydrogen storage density, the 2 heats of combustion of hydrogen and other combustibles in the fuel system, and the cooling capacity of the fuel and the storage system are important parameters in the evaluation of candidate fuel systems. It should also be realized that for any improvement in hydrogen storage density a certain penalty has to be paid in terms of increased mass, decreased specific impulse, or increased cost and complexity of tankage, fuel feed systems and technology development. Increase in fuel mass may be at least partially compensated by decreases in some mass components such as the tankage mass and thermal protection mass. Decrease in the specific impulse may be offset by a decrease in drag losses so that the effective specific impulse may not be reduced as much as the specific impulse. These effects depend on, among others, the flight trajectory, whether the plane is designed as a launch vehicle or as a hypersonic transport plane, the structural design, the type of engines to be used, and the switchover Mach numbers for the engines. Only the first two effects are considered in the present work. To account for them, differences in the effective specific impulses and the payload capacities are taken to be the additional evaluation criteria. In order to quantify the basis for the evaluation of the candidate fuel systems they are first thermodynamically screened with respect to their hydrogen density, energy content, and cooling capacities. The most promising systems are then evaluated using the developed model in terms of the payload capacities and effective specific impulses of the corresponding vehicles. The results of the ther- modynamic evaluation are given in Appendix A, the flow chart for the computer model is presented in Appendix B, the computer program is included in Appendix C, information on the computer program is in Appendix D, and the results of some computations are included in Appendix E. 3 MODELINGCONSIDERATIONS At the start of this study the only tools available to us for the comparison of the performances of vehicles with different structures and engines were the ongoing work by Dorrington I on alpha-beta relationships and the ASP computer program developed at NASA Langley Research Center 2 for the assessment of the effects of component size changes on the aircraft performance. The former was still!under development and the latter could only be applied to vehicles with turbojet - ramjet engines using liquid hydrogen or methane and was limited to Mach numbers less than 4.5. Recently, we became aware of a similar study done for the Air Force Wright Research and Development Center by Aerojet TechSystems and Boeing Aerospace, which used inhouse codes for engine analysis and tra- jectory optimization and compared the performances of vehicles using a variety of fuels based on ammonia and boron hydrides to the performance of a vehicle using slush hydrogen. The dissemination of this information was restricted and therefore, it did not have much influence on the present study. Since the design of NASP is not finalized yet, there is no need to accurately predict the performance of a certain hypothetical vehicle. The purpose of this study will be better served by a simple model which can compare the payload capacity and effective specific impulse of various vehicles to those of a vehicle using liquid hydrogen. To achieve this we used the conceptual design approach of chemical process design. Accordingly we started with the simplest possible model and added details and complexities step by step until the model produced sufficient information of acceptable accuracy. The starting model only considered the hypersonic air-breathing phase of the flight. The reasoning was that during the subsonic - supersonic phase all the vehicles could use identical engines and fuels if they had the same gross lift off mass. The mass change was calculated by a macroscopic energy balance similar to the approach used by Jones and Donaldson 3, which used a specified thrust 4 to drag ratio to account for the drag losses. The initial thrust to drag ratio was assumed to be the same for all vehicles and it is used to determine the engine size which is assumed to be fixed. The thrust and drag (and, hence, their ratio) were allowed to change during the hypersonic flight. The specific impulse was calculated at the beginning and at the end of the hypersonic phase. The details of the engine was not considered. The combustion chamber pressure was determined by specifying a compression ratio. Products of constant pressure combustion at equilibrium was determined by the chemical equilibrium code developed by NASALewis Research Center. Exit velocities were computed by assuming frozen expansion in an ideal nozzle to ambient pressure. All vehicles were assumed to have the same gross lift off mass and any weight penalty manifested itself as reduced payload capacity. This enabled several mass components such as the thrust structure mass and the engine mass which are functions of the initial vehicle mass to be the same for all vehicles and simplified the analysis considerably. This initial model failed to discriminate effectively between vehicles with different fuel systems due to its various shortcomings. The problems and the way they are dealt with in the final program are summarized below: i. If chemical binding is used to increase hydrogen storage density, the extra mass introduced should replace an equal amount of mass that already exists on board the vehicle to prevent a reduction in the payload capacity. One method which seemed feasible was the possibility of using the extra mass as the rocket fuel for the final stage of flight after the extraction of hydrogen to be used as the air breathing phase fuel. For this purpose a section was added to the program to evaluate the performance of the rocket phase and compute the fuel requirements. This phase of the flight was assumed to be free of drag losses. Instead of assuming a specific impulse and calculating the mass ratio using this specific impulse, a macroscopic energy balance was used to obtain the mass 5 ratio and the specific impulse was then calculated using this information. This was done to account for different specific impulses of different fuels. 2. Since the vehicle sizes will be different due to different fuel volumes, the drag encountered by each vehicle will be different. In order to account for this the hypersonic flight phase was investigated in more detail. The initial simple model was used to find the mass ratio for the subsonic-supersonic phase. Since the same average trust to drag ratio is used for all vehicles, the vehicles with smaller drag will have a larger thrust during this first phase. This will affect the required engine and thrust structure masses. At this level of sophistication of the model this effect is ignored. It was also assumed that the vehicles to be compared will have the samethrust to drag ratio at the commencement of the hypersonic flight phase. From this information the thrust and the capture area for each vehicle is obtained and assumed to be fixed for the entire flight. The effective specific impulse is computed at each i00 m altitude step and after every i0 steps the differential equation giving the mass ratio for the interval was integrated numerically. Since there is no dataagainst which the results of the proposed model can be checked, the ability of the model to represent the performance of a hypersonic vehicle can be verified only by checking if the magnitude and the variation of quantities such as thrust, drag, and specific impulse are technically reasonable. The introduction of details mentioned above also provided information which were used for this purpose. 3. The original model used a specified compression ratio to calculate the combustion chamber pressure and could not account for the effect of Mach number on forebody compression. In the final model a more realistic approach is used. The forebody and engine geometry used is taken from Ikawa 4 and his method is used to obtain the conditions at the combustion chamber exit. For the aftbody expansion we kept the simplifying assumption of isentropic, frozen expansion. 4. In the initial model the drag coefficient for the vehicle was taken to be dependent only on the angle of attack. Since the omission of the Mach number dependence produced unsatisfactory drag and thrust profiles, the equation used to obtain the drag coefficients is modified to include Mach number dependence. This is done by fitting an equation to the curve given by Dorrington 2. 5. The initial model was modified to allow the specification of varying dynamic pressures and angle of attack values during the hypersonic flight phase. At the present, three different values can be specified at three selected flight Mach numbers. The program can easily be changed to increase this number. The flow chart for the final model is given in Appendix B and the program listing in Appendix C. RESULTS The results for some potential fuel systems are summarized in Table I. The entries in this table are the differences from the corresponding values for a vehicle using liquid hydrogen as fuel for the entire flight. These results were obtained for a fixed set of conditions given below: Dynamic pressure = 47882 Pa (i000 ibf/ft 2) Gross lift-off mass = 300,000 kg Orbital altitude = 200,000 m Orbital velocity = 8030 m/s Angle of attack = 2 degrees Switchover Mach number for hypersonic propulsion = 3 switchover Mach number for rocket propulsion = 12 7 Table I. Comparison of hypersonic vehicles using various fuel systems with the vehicle using liquid hydrogen. Negative sign indicates a value lower than that of the LH2 vehicle. Difference Difference in Difference Approximate Difference in F_el System effective specific First Second Third in vehicle dry mass (%) in payload difference pfhuaesle pfhuaesle pfhuaesle volume (%) c(a%p)acity in GLOW (%) irmapnugles)e (%()Mach 3-12 SH2 SH2 SH2 - 11 3.2 + 7.2 1.8 + 1.2 to + 2.9 CH4 CH4 CH4 - 50 18 80 + 20 - 53 to - 66 CH4-H2 CH4-H2 CH4-H2 - 12 3.8 + 0.6 0.14 - 3.6 to - 3.2 CH4 H2 CH4 - 48 15 55 + 13 + 5.4 to + 13 CH4 NH3 CH4 - 48 19 89 + 22 - 76 to - 53 co C3H8 NH3B5H9 C3H8 - 56 20 89 + 22 - 69 to - 63 CH4 H2 CO CO cannot be used alone as the rocket phase fuel. LH2 H2(C6H12) C6H6 8.6 - 11 83 + 20 + 0.9 to + 2.3 LH2 H2(C7H14) C7H8 8.0 - 11 83 + 20 + 0.8 to + 2.1 LH2 H2(B2H6) 8 16 9.1 50 + 12 + 1.7 to + 4.2 LH2 H2(ALH3) AL - 34 - 16 96 + 24 + 3.7 to + 9.3 LH2 H2(LIH) LI - 17 - 11 - 75 + 18 + 1.8 to + 4.5 LH2 H2(NH3B5H9) BN - 19 - 19 - 129 + 32 + 2.0 to + 5.1 LH2 H2(NH3810H13) BN - 2.7 16 - 131 + 32 + 0.3 to + 0.8 LH2 H2(N2H5B5H9) BN - 19 18 129 + 32 + 2.0 to + 5.1