Gaseous oxygen/gaseous hydrogen injector element modeling
Gaseous oxygen/hydrogen injector element modeling based on composition profile measurements for cold flows
Calhoon, D. F., Sr.↗
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Effect of combustor parameters on the stability of gaseous hydrogen-liquid oxygen engine
Combustor parameters in gaseous hydrogen-liquid oxygen engine
Feiler, C. E.↗
Detonation of cryogenic gaseous hydrogen-oxygen mixtures
The accidental mixing and detonation of oxygen-hydrogen mixtures is a serious aerospace industry hazard. The detonation characteristics of cryogenic mixtures of gaseous hydrogen and oxygen have accordingly been studied at various initial pressures and equivalence ratios at 100 K, and the results compared with numerical computations of idealized, steady-state, one-dimensional Zeldovich-von Neumann-Doering detonation structure with detailed chemical reaction kinetics. The predictions thus obtained for critical tube diameter, on the basis of calculated reaction-zone thickness, are found to agree reasonably well with experimental data.
Plaster, M.↗
Cold-flow optimization of gaseous oxygen/gaseous hydrogen injectors for the Space Shuttle APS thruster
Gaseous oxygen/gaseous hydrogen auxiliary propulsion engines, considering cold flow experiments with nonreactive simulant gases
Clapp, S. D.↗
A 25-LBF gaseous oxygen/gaseous hydrogen thruster for space station application
A prototype 25 lb sub f gaseous oxygen/gaseous hydrogen thruster for Space Station propulsion application was designed and fabricated by Rocketdyne and endurance tested at the NASA/Marshall space Flight Center. The thruster incorporates a regeneratively cooled thrust chamber with a nozzle exit area ratio of 30, a 12-element coaxial injector, a spark igniter, and close-coupled propellant valves. Test results indicate that all major technology issues for long-life gaseous oxygen/gaseous hydrogen thrusters for Space Station application have been resolved.
Heckert, B. J.↗
Hydrogen adsorption and diffusion, and subcritical-crack growth in high-strength steels and nickel base alloys
Coordinated studies of the kinetics of crack growth and of hydrogen adsorption and diffusion were initiated to develop information that is needed for a clearer determination of the rate controlling process and possible mechanism for hydrogen enhanced crack growth, and for estimating behavior over a range of temperatures and pressures. Inconel 718 alloy and 18Ni(200) maraging steel were selected for these studies. 18Ni(250) maraging steel, 316 stainless steel, and iron single crystal of (111) orientation were also included in the chemistry studies. Crack growth data on 18Ni(250) maraging steel from another program are included for comparison. No sustained-load crack growth was observed for the Inconel 718 alloy in gaseous hydrogen. Gaseous hydrogen assisted crack growth in the 18Ni maraging steels were characterized by K-independent (Stage 2) extension over a wide range of hydrogen pressures (86 to 2000 torr or 12 kN/m2 to 266 kN/m2) and test temperatures (-60 C to +100 C). The higher strength 18Ni(250) maraging steel was more susceptible than the lower strength 200 grade. A transition temperature was observed, above which crack growth rates became diminishingly small.
Wei, R. P.↗
OH-radical imaging and velocity field measurements in a gaseous hydrogen/oxygen rocket
Detailed measurements of the OH-radical and velocity fields for a shear coaxial unielement rocket chamber burning gaseous hydrogen and gaseous oxygen propellants are reported. The velocity measurements indicate that a highly unsteady flowfield results from the turbulent combustion of the coaxial jets. The core of the oxygen stream is observed to remain intact well downstream of the injector face while the maximum turbulent intensities in the shear layer region between the hydrogen and oxygen flows were measured to be about 25 percent. Both 1D and 2D laser-induced fluorescence studies of the OH-radical are reported and provide information on the evolving flame structure under rocketlike conditions. Both single and multiple laser pulse images indicate that the structure of the unsteady flowfield is directly reflected in the combustion zone flame structure. Preliminary results regarding analysis approaches and interpretation of the data are also presented.
Moser, M. D.↗
The water-cryogen heat exchanger
Heat exchanger, using water as heat medium, converts liquid hydrogen to gaseous hydrogen at a very high rate. Possible applications include treatment of liquified natural gas in cities to bring the gas on-line quickly, conversion of liquid oxygen and liquid nitrogen for steel mills, and high volume inert purging.
Bartlit, J. R.↗
High Pressure Reverse Flow APS Engine
A design and test demonstration effort was undertaken to evaluate the concept of the reverse flow engine for the APS engine application. The 1500 lb (6672 N) thrust engine was designed to operate on gaseous hydrogen and gaseous oxygen propellants at a mixture ratio of 4 and to achieve the objective performance of 435 sec (4266 Nsec/kg) specific impulse. Superimposed durability requirements called for a million-cycle capability with 50 hours duration. The program was undertaken as a series of tasks including the initial preliminary design, design of critical test components and finally, the design and demonstration of an altitude engine which could be used interchangeably to examine operating parameters as well as to demonstrate the capability of the concept. The program results are reported with data to indicate that all of the program objectives were met or exceeded within the course of testing on the program. The analysis effort undertaken is also reported in detail and supplemented with test data in some cases where prior definitions could not be made. The results are contained of these analyses as well as the test results conducted throughout the course of the program. Finally, the test data and analytical results were combined to allow recommendations for a flight weight design. This preliminary design effort is also detailed.
Senneff, J. M.↗
Investigation of gaseous propellant combustion and associated injector/chamber design guidelines
Injector design criteria are provided for gaseous hydrogen-gaseous oxygen propellants. Design equations and procedures are presented which will allow an injector-chamber designer to a priori estimate of the performance, compatibility and stability characteristics of prototype injectors. The effects of chamber length, element geometry, thrust per element, mixture ratio, impingement angle, and element spacing were evaluated for four element concepts and their derivatives. The data from this series of tests were reduced to a single valued mixing function that describes the mixing potential of the various elements. Performance, heat transfer and stability data were generated for various mixture ratios, propellant temperatures, chamber pressures, contraction ratios, and chamber lengths. Applications of the models resulted in the design of procedures, whereby the performance and chamber heat flux can be calculated directly, and the injector stability estimated in conjunction with existing models.
Calhoon, D. F.↗
Measurement of the turbulence in a gaseous rocket combustion chamber
The intensity of turbulence and the Lagrangian correlation coefficient in a gaseous hydrogen (GH2)-gaseous oxygen (GOX) rocket combustion chamber are studied. A helium tracer gas was injected at an upstream point on the combustion chamber centerline. Samples collected downstream were analyzed for the spreading of the tracer gas concentration. Using Taylor's turbulent diffusion theory and Gosman's numerical method, turbulence parameters were found. In order to evaluate turbulence field inhomogeneity in the axial direction, Taylor's theory was extended. An exponential function was used to approximate the Lagrangian correlation coefficient. It was found that turbulence intensity varies from 15% near the injector to 4% at the nozzle entrance, and that the Lagrangian coefficient rapidly decreases when the dispersion time difference increases.
Tou, P. P.↗
Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle
Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1030:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.
Smith, Tamara A.↗
Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle
Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1039:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.
Smith, Tamara A.↗
Ablative material testing for low-pressure, low-cost rocket engines
The results of an experimental evaluation of ablative materials suitable for the production of light weight, low cost rocket engine combustion chambers and nozzles are presented. Ten individual specimens of four different compositions of silica cloth-reinforced phenolic resin materials were evaluated for comparative erosion in a subscale rocket engine combustion chamber. Gaseous hydrogen and gaseous oxygen were used as propellants, operating at a nominal chamber pressure of 1138 kPa (165 psi) and a nominal mixture ratio (O/F) of 3.3. These conditions were used to thermally simulate operation with RP-1 and liquid oxygen, and achieved a specimen throat gas temperature of approximately 2456 K (4420 R). Two high-density composition materials exhibited high erosion resistance, while two low-density compositions exhibited approximately 6-75 times lower average erosion resistance. The results compare favorably with previous testing by NASA and provide adequate data for selection of ablatives for low pressure, low cost rocket engines.
Richter, G. Paul↗
Rocket-in-a-Duct Performance Analysis
An axisymmetric, 110 N class, rocket configured with a free expansion between the rocket nozzle and a surrounding duct was tested in an altitude simulation facility. The propellants were gaseous hydrogen and gaseous oxygen and the hardware consisted of a heat sink type copper rocket firing through copper ducts of various diameters and lengths. A secondary flow of nitrogen was introduced at the blind end of the duct to mix with the primary rocket mass flow in the duct. This flow was in the range of 0 to 10% of the primary massflow and its effect on nozzle performance was measured. The random measurement errors on thrust and massflow were within +/-1%. One dimensional equilibrium calculations were used to establish the possible theoretical performance of these rocket-in-a-duct nozzles. Although the scale of these tests was small, they simulated the relevant flow expansion physics at a modest experimental cost. Test results indicated that lower performance was obtained at higher free expansion area ratios and longer ducts, while, higher performance was obtained with the addition of secondary flow. There was a discernable peak in specific impulse efficiency at 4% secondary flow. The small scale of these tests resulted in low performance efficiencies, but prior numerical modeling of larger rocket-in-a-duct engines predicted performance that was comparable to that of optimized rocket nozzles. This remains to be proven in large-scale, rocket-in-a-duct tests.
Schneider, Steven J.↗
Catalyzed Ignition of Bipropellants in Microtubes
This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with attitude control estimated to be in the 10 milli-Newton thrust class. These combustors are anticipated to be manufactured using Micro Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the quenching diameter of the propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in geometry can be used in fundamental simulations to more carefully characterize the measured heat transfer and pressure losses for validation purposes. Experimentally, we investigate the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter micro-tubes, with special emphases on ignition and extinction processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame thickness and reaction zone thickness calculations predict that the diameters of our test apparatus are below the quenching diameter of the propellants in sub-atmospheric tests. Temperature and pressure rises in resistively heated platinum and palladium micro-tubes are used as an indication of exothermic reactions. Specific data on mass flow versus preheat temperature required to achieve ignition are presented.
Schneider, Steven J.↗
Catalyzed Combustion of Bipropellants for Micro-Spacecraft Propulsion
This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with attitude control estimated to be in the 10 milli-Newton thrust class. These combustors are anticipated to be manufactured using Micro Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the quenching diameter of the propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in geometry can be used in fundamental simulations for validation purposes. Experimentally, we investigated the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter microtubes, with special emphases on ignition processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame thickness and reaction zone thickness calculations predict that the diameters of our test apparatus are below the quenching diameter of the propellants in sub-atmospheric tests. Temperature and pressure rise in resistively heated platinum and palladium microtubes was used as an indication of exothermic reactions. Specific data on mass flow versus preheat temperature required to achieve ignition are presented. With a plug flow model, the experimental conditions were simulated with detailed gas-phase chemistry, thermodynamic properties, and surface kinetics. Computational results generally support the experimental findings, but suggest an experimental mapping of the exit temperature and composition is needed.
Schneider, Steven J.↗