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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

A parametric study of a gas-generator airturbo ramjet (ATR)

Parametric engine performance calculations were carried out for an airturbo ramjet (ATR). A LOX-LH2 rocket powered turbine powered the compressor. The engine was flown over a typical flight path up to Mach 5 to show the effect of engine off design operation. The compressor design efficiency, compressor pressure ratio, rocket turbine efficiency, rocket turbine inlet temperature, and rocket chamber pressure were varied to show their effect on engine net thrust and specific impulse at Mach 5 cruise. Estimates of engine weights as a function of the ratio of compressor air to rocket propellant flow and rocket chamber pressure are also included. In general, the Mach 5 results indicate that increasing the amount of rocket gas produced increased thrust but decreased the specific impulse. The engine performance was fairly sensitive to rocket chamber pressure, especially at higher compressor pressure ratios. At higher compressor pressure ratios, the engine thrust was sensitive to turbine inlet temperature. At all compressor pressure ratios, the engine performance was not sensitive to compressor or turbine efficiency.

Snyder, Christopher A.↗
NASA NTRS2005

Benchmark Wall Heat Flux Data for a GO2/GH2 Single Element Combustor

Wall heat flux measurements in a 1.5 in. diameter circular cross-section rocket chamber for a uni-element shear coaxial injector element operating on gaseous oxygen (GOz)/gaseous hydrogen (GH,) propellants are presented. The wall heat flux measurements were made using arrays of Gardon type heat flux gauges and coaxial thermocouple instrumentation. Wall heat flux measurements were made for two cases. For the first case, GOZ/GHz oxidizer-rich (O/F=l65) and fuel-rich preburners (O/F=1.09) integrated with the main chamber were utilized to provide vitiated hot fuel and oxidizer to the study shear coaxial injector element. For the second case, the preburners were removed and ambient temperature gaseous oxygen/gaseous hydrogen propellants were supplied to the study injector. Experiments were conducted at four chamber pressures of 750, 600, 450 and 300psia for each case. The overall mixture ratio for the preburner case was 6.6, whereas for the ambient propellant case, the mixture ratio was 6.0. Total propellant flow was nominally 0.27-0.29 Ibm/s for the 750 psia case with flowrates scaled down linearly for lower chamber pressures. The axial heat flux profile results for both the preburner and ambient propellant cases show peak heat flux levels a t axial locations between 2.0 and 3.0 in. from the injector face. The maximum heat flux level was about two times greater for the preburner case. This is attributed to the higher injector fuel-to-oxidizer momentum flux ratio that promotes mixing and higher initial propellant temperature for the preburner case which results in a shorter reaction zone. The axial heat flux profiles were also scaled with respect to the chamber pressure to the power 0.8. The results at the four chamber pressures for both cases collapsed to a single profile indicating that at least to first approximation, the basic fluid dynamic structures in the flow field are pressure independent as long as the chamber/njector/nozzle geometry and injection velocities remain the same.

Marshall, William M.↗
NASA NTRS1962

An Experimental Investigation of the Performance of the Nitrogen Tetroxide-Hydrazine System in the Oxidizer-Rich and Fuel-Rich Regions

The results of an experimental program directed toward determining some of the operational characteristics of the nitrogen tetroxide-hydrazine propellant system under oxidizer-rich and fuel-rich conditions are reported. Data are presented for the mixture-ratio ranges of 0 to 0.55 at a nominal chamber pressure of 300 psia and characteristic chamber length of 250 in., and at mixture ratios of 5.9 to 13.0 at nominal chamber pressures of 400 and 500 psia and characteristic chamber lengths of 100 and 3900 in. The relationship of the actual performance data obtained in each region with those predicted from thermochemical performance calculations is presented. From this comparison, it is concluded that in neither region are equilibrium conditions obtained and that utilization of performance data obtained from assumptions of equilibrium will lead to serious errors. In the oxidizer-rich region investigated, temperatures considerably below those predicted occur throughout the range of mixture ratios investigated because of the lack of exothermic dissociation of the nitrogen oxides. In the fuel-rich region, below a mixture ratio of 0.3, temperatures considerably higher than those predicted occur because of the lack of endothermic dissociation of ammonia.

Chilenski, J. J.↗
NASA NTRS1987

Wall relaxation in growing stems: comparison of four species and assessment of measurement techniques

This study was carried out to develop improved methods for measuring in-vivo stress relaxation of growing tissues and to compare relaxation in the stems of four different species. When water uptake by growing tissue is prevented, in-vivo stress relaxation occurs because continued wall loosening reduces wall stress and cell turgor pressure. With this procedure one may measure the yield threshold for growth (Y), the turgor pressure in excess of the yield threshold (P-Y), and the physiological wall extensibility (phi). Three relaxation techniques proved useful: "turgor-relaxation", "balance-pressure" and "pressure-block". In the turgor-relaxation method, water is withheld from growing tissue and the reduction in turgor is measured directly with the pressure probe. This technique gives absolute values for P and Y, but requires tissue excision. In the balance-pressure technique, the excised growing region is sealed in a pressure chamber, and the subsequent reduction in water potential is measured as the applied pressure needed to return xylem sap to the cut surface. This method is simple, but only measures (P-Y), not the individual values of P and Y. In the pressure-block technique, the growing tissue is sealed into a pressure chamber, growth is monitored continuously, and just sufficient pressure is applied to the chamber to block growth. The method gives high-resolution kinetics of relaxation and does not require tissue excision, but only measures (P-Y). The three methods gave similar results when applied to the growing stems of pea (Pisum sativum L.), cucumber (Cucumis sativus L.), soybean (Glycine max (L.) Merr.) and zucchini (Curcubita pepo L.) seedlings. Values for (P-Y) averaged between 1.4 and 2.7 bar, depending on species. Yield thresholds averaged between 1.3 and 3.0 bar. Compared with the other methods, relaxation by pressure-block was faster and exhibited dynamic changes in wall-yielding properties. The two pressure-chamber methods were also used to measure the internal water-potential gradient (between the xylem and the epidermis) which drives water uptake for growth. For the four species it was small, between 0.3 and 0.6 bar, and so did not limit growth substantially.

NASA Program Space Biology↗
NASA NTRS2014

Combustion Stability Characteristics of the Project Morpheus Liquid Oxygen / Liquid Methane Main Engine

The project Morpheus liquid oxygen (LOX) / liquid methane (LCH4) main engine is a Johnson Space Center (JSC) designed ~5,000 lbf-thrust, 4:1 throttling, pressure-fed cryogenic engine using an impinging element injector design. The engine met or exceeded all performance requirements without experiencing any in- ight failures, but the engine exhibited acoustic-coupled combustion instabilities during sea-level ground-based testing. First tangential (1T), rst radial (1R), 1T1R, and higher order modes were triggered by conditions during the Morpheus vehicle derived low chamber pressure startup sequence. The instability was never observed to initiate during mainstage, even at low power levels. Ground-interaction acoustics aggravated the instability in vehicle tests. Analysis of more than 200 hot re tests on the Morpheus vehicle and Stennis Space Center (SSC) test stand showed a relationship between ignition stability and injector/chamber pressure. The instability had the distinct characteristic of initiating at high relative injection pressure drop at low chamber pressure during the start sequence. Data analysis suggests that the two-phase density during engine start results in a high injection velocity, possibly triggering the instabilities predicted by the Hewitt stability curves. Engine ignition instability was successfully mitigated via a higher-chamber pressure start sequence (e.g., ~50% power level vs ~30%) and operational propellant start temperature limits that maintained \cold LOX" and \warm methane" at the engine inlet. The main engine successfully demonstrated 4:1 throttling without chugging during mainstage, but chug instabilities were observed during some engine shutdown sequences at low injector pressure drop, especially during vehicle landing.

Melcher, John C.↗
NASA NTRS1995

Plume Diagnostics of the RSRM Static Firings for the Pressure Perturbation Studies

During the STS-54 launch (RSRM-29), the right hand solid rocket motor experienced a 13.9 psi chamber pressure perturbation at 67 seconds into the motor operation. This pressure augmentation equated to a thrust change of 51 klb. Concerns were raised regarding the adverse effects of this thrust imbalance on the shuttle system and the overall thrust into the external tank structural elements. Pressure perturbations have been observed in solid rocket motors due to expulsion of igniter or insulation materials; the motor thrust during such events drop abruptly before rising. However, the RSRM motors do not exhibit such behavior during the large chamber pressure perturbation events. Several scenarios were investigated to explain these pressure perturbations in the RSRM motors based on a fault tree developed after STS-54. Of these, the expulsion of the slag accumulated in the submerged nozzle region appeared to be the most plausible scenario to explain the observations. Slag is a natural combustion product of aluminized solid rocket motors. The RSRM propellant contains 16% by weight of aluminum. Any ejection of this slag mass during nozzle vectoring or other side loads on the motor will result in the chamber pressure perturbation. Two RSRM static firings were instrumented extensively to further understand the slag expulsion phenomenon in the RSRM and the associated pressure perturbations.

Mathias, Edward C.↗
NASA NTRS1991

Subscale LOX/hydrogen testing with a modular chamber and a swirl coaxial injector

To support basic injector and chamber technology, the response of a modular calorimeter chamber to an injector with swirl coaxial elements was investigated. The tests supported both the Advanced Launch System (ALS) and the Space Shuttle Main Engine programs. Original test plans included chamber pressures up to 2250 psia, the current ALS engine criteria, but hardware limitations forced conditions to be reduced. With four different chamber configurations, a variety of data was obtained with chamber pressures ranging from 1483 psia to 1679 psia and mixture ratios from 5.24 to 6.90. The swirl coaxial injector showed good chamber wall compatibility, despite misaligned elements, and its high performance was independent of fuel. Heating rates and wall temperatures were acceptable and close to predict profiles, and after scaling to a chamber pressure of 2250 psia, heating rates remained acceptable. In addition, a chamber spool made with the vacuum plasma spray process survived hot-fire testing with no detrimental effects.

Elam, Sandra K.↗
NASA NTRS1997

High Pressure, Earth-storable Rocket Technology

The effect of elevated chamber pressure on combustion efficiency and heat transfer has been determined at the 100 lbf (445 N) thrust level for nitrogen tetroxide propellants. Measurements were made up to 500 psia (3.45 Mpa) with testbed hardware; tests at 100 psia (0.690 MPa) and 250 psia (1.72 MPa) were made with radiation-cooled rhenium chambers. The first task of the program served to determine desirable thruster applications and operating conditions: high total impulse, i.e. communication satellite or spacecraft bus axial engines, at chamber pressures up to 250 psia (1.72 MPa) pressure-fed, or up to 500 psia (3.45 MPa) pump-fed. The hardware modifications and testing required to obtain the data were determined in Task 2, which included design-support hot fire tests; supplemental hardware, including a 250 psia (1.72 MPa) Pc rhenium chamber and a 20% fuel-film cooled platelet injector was fabricated in Task 3. Testing showed that satisfactory operation of Ir-Re radiation chambers is assured at pressures up to 250 psia and may be possible up to 500. The heat transfer data obtained show good correlation with throat Reynolds number and are generally under values given by the simplified Bartz equation; chambers equilibrium temperatures match predicted values. Preliminary optimization of trip configuration and mixture ratio were made; Isp performance from thrust measurements was within 1% of predicted values. Stability, compatibility, and front-end thermal management were determined to be satisfactory.

Jassowski, D. M.↗
NASA NTRS1997

High Pressure, Earth-storable Rocket Technology

The effect of elevated chamber pressure on combustion efficiency and heat transfer has been determined at the 100 lbf (445 N) thrust level for nitrogen tetroxide propellants. Measurements were made up to 500 psia (3.45 MPa) with testbed hardware; tests at 100 psia (0.690 MPa) and 250 psia (1.72 MPa) were made with radiation-cooled rhenium chambers. The first task of the program served to determine desirable thruster applications and operating conditions: high total impulse, i.e., communication satellite or spacecraft bus axial engines, at chamber pressures up to 250 psia (1.72 MPa) pressure-fed, or up to 500 psia (3.45 MPa) pump-fed. The hardware modifications and testing required to obtain the data were determined in Task 2, which included design-support hot fire tests; supplemental hardware, including a 250 psia (1.72 MPa) Pc rhenium chamber and a 20% fuel-film cooled platelet injector was fabricated in Task 3. Testing showed that satisfactory operation of Ir-Re radiation chambers is assured at pressures up to 250 psia and may be possible up to 500. The heat transfer data obtained show good correlation with throat Reynolds number and are generally under values given by the simplified Bartz equation; chambers equilibrium temperatures match predicted values. Preliminary optimization of trip configuration and mixture ratio were made; Isp performance from thrust measurements was within 1% of predicted values. Stability, compatibility, and front-end thermal management were determined to be satisfactory.

Jassowski, D. M.↗
NASA NTRS1973

Pump for delivering heated fluids

A thermomechanical pump particularly suited for use in pumping a warming fluid obtained from an RTG (Radioisotope Thermal Generator) through science and flight instrumentation aboard operative spacecraft is described. The invention is characterized by a pair of operatively related cylinders, each including a reciprocating piston head dividing the cylinder into a pressure chamber confining therein a vaporizable fluid, and a pumping chamber for propelling the warming fluid, and a fluid delivery circuit for alternately delivering the warming fluid from the RTG through the pressure chamber of one cylinder to the pumping chamber of the other cylinder, whereby the vaporizable fluid within the pair of pressure chambers alternately is vaporized and condensed for driving the associated pistons in pumping and intake strokes.

Sabelman, E. E.↗
NASA NTRS1992

Plant growth modeling at the JSC variable pressure growth chamber - An application of experimental design

This paper describes the approach and results of an effort to characterize plant growth under various environmental conditions at the Johnson Space Center variable pressure growth chamber. Using a field of applied mathematics and statistics known as design of experiments (DOE), we developed a test plan for varying environmental parameters during a lettuce growth experiment. The test plan was developed using a Box-Behnken approach to DOE. As a result of the experimental runs, we have developed empirical models of both the transpiration process and carbon dioxide assimilation for Waldman's Green lettuce over specified ranges of environmental parameters including carbon dioxide concentration, light intensity, dew-point temperature, and air velocity. This model also predicts transpiration and carbon dioxide assimilation for different ages of the plant canopy.

Miller, Adam M.↗
NASA NTRS1996

Design of Plant Gas Exchange Experiments in a Variable Pressure Growth Chamber

Sustainable human presence in extreme environments such as lunar and martian bases will require bioregenerative components to human life support systems where plants are used for generation of oxygen, food, and water. Reduced atmospheric pressures will be used to minimize mass and engineering requirements. Few studies have assessed the metabolic and developmental responses of plants to reduced pressure and varied oxygen atmospheres. The first tests of hypobaric pressures on plant gas exchange and biomass production at the Johnson Space Center will be initiated in January 1996 in the Variable Pressure Growth Chamber (VPGC), a large, closed plant growth chamber rated for 10.2 psi. Experiments were designed and protocols detailed for two complete growouts each of lettuce and wheat to generate a general database for human life support requirements and to answer questions about plant growth processes in reduced pressure and varied oxygen environments. The central objective of crop growth studies in the VPGC is to determine the influence of reduced pressure and reduced oxygen on the rates of photosynthesis, dark respiration, evapotranspiration and biomass production of lettuce and wheat. Due to the constraint of one experimental unit, internal controls, called pressure transients, will be used to evaluate rates of CO2 uptake, O2 evolution, and H2O generation. Pressure transients will give interpretive power to the results of repeated growouts at both reduced and ambient pressures. Other experiments involve the generation of response functions to partial pressures of O2 and CO2 and to light intensity. Protocol for determining and calculating rates of gas exchange have been detailed. In order to build these databases and implement the necessary treatment combinations in short time periods, specific requirements for gas injections and removals have been defined. A set of system capability checks will include determination of leakage rates conducted prior to the actual crop growouts. Schedules of experimental events for lettuce and wheat are outlined and include replications in time of diurnal routines, pressure transients, variable pO2, pO2/pCO2 ratio, and light intensity responses.

Corey, Kenneth A.↗
NASA NTRS2002

Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio (LD). This incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L/D vortex chamber concept for gel propellants was conducted by Michaels. He used both triplet (two oxidizer orifices and one fuel orifice) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called the subject injection scheme an Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 9295, can be obtained. MSFC and the U. S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen/hydrocarbon propellant system. This vortex chamber concept is currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept for the liquid oxygen (LOX) hydrocarbon fuel (RP-1) system has been derived from the one for the gel propellant. An unlike impinging injector was employed to deliver the propellants to the chamber. MSFC is also conducting an alternative injection scheme, called the chasing injector, associated with this vortex chamber concept. In this injection technique, both propellant jets and their impingement point are in the same chamber cross-sectional plane. Long duration tests (approximately up to 15 seconds) will be conducted on the ISVC to study the thermal effects. This paper will report the progress of the subject efforts at NASA Marshall Space Flight Center. Thrust chamber performance and thermal wall compatibility will be evaluated. The chamber pressures, wall temperatures, and thrust will be measured as appropriate. The test data will be used to validate CFD models, which, in turn, will be used to design the optimum vortex chambers. Measurements in the previous tests showed that the chamber pressures vary significantly with radius. This is due to the existence of the vortices in the chamber flow field. Hence, the combustion efficiency may not be easily determined from chamber pressure. For this project, measured thrust data will be collected. The performance comparison will be in terms of specific impulse efficiencies. In addition to the thrust measurements, several pressure and temperature readings at various locations on the chamber head faceplate and the chamber wall will be made. The first injector and chamber were designed and fabricated based on the available data and experience gained during gel propellant system tests by the U.S. Army. The alternate injector for the ISVC was also fabricated. Hot-fire tests of the vortex chamber are about to start and are expected to complete in February of 2003 at the TS115 facility of MSFC.

Trinh, Huu P.↗
NASA NTRS1988

Orbital transfer vehicle 3000 LBF thrust chamber assembly hot fire test program

The Aerojet Orbital Transfer Vehicle (OTV) Thrust Chamber Assembly (TCA) concept consists of a hydrogen cooled chamber, and annular injector, and an oxygen cooled centerbody. The hot fire testing of a heat sink version of the chamber with only the throat section using hydrogen cooling is documented. Hydraulic performance of the injector and cooled throat were verified by water flow testing prior to TCA assembly. The cooled throat was proof tested to 3000 psia to verify the integrity of the codeposited EF nickel-cobalt closeout. The first set of hot fire tests were conducted with a heat sink throat to obtain heat flux information. After demonstration of acceptable heat fluxes, the heat sink throat was replaced with the LH2 cooled throat section. Fourteen tests were conducted with a heat sink chamber and throat at chamber pressures of 85 to 359 psia. The injector face was modified at this time to add more face coolant flow. Ten tests were then conducted at chamber pressures of 197 to 620 psia. Actual heat fluxes at the higher chamber pressure range were 23 percent higher than the average of 10 Btu/in 2 predicted.

Schneider, Judy↗
NASA NTRS1992

Analysis of the lettuce data from the variable pressure growth chamber at NASA Johnson Space Center: A three-stage nested design model

A model of three-stage nested experimental design was applied to analyze the lettuce data obtained from the variable pressure growth chamber test bed at NASA-Johnson Space Center. From the results of an application of the analysis of variance and covariance on the data set, it was noted that all of the (uncontrollable) factors, Side, Zone, Height and (controllable) PAR (photosynthetically active radiation), had nonhomogeneous effects on the dry weight of the edible biomass of lettuce per pot. Incidentally, the variations accountable to the (uncontrollable) factorial heterogeneities are merely 9 percent and 17 percent of the total variation for both the first and second crop test, respectively. After adjusting for the PAR as a covariate in the no-intercept model, the accountable variations to all the four factors are 94 percent and 92 percent for the first and the second crop test, respectively. With the use of a no-intercept simple linear regression model, the accountable variations to the factor PAR are 92 percent and 90 percent for the first and the second crop test, respectively. Evidently, the (controllable) factor PAR is the dominating one.

Lee, Tze-San↗
NASA NTRS1983

Advanced LOX/H sub 2 engine technologies for future OTVs

The optimization methods and trade studies which were used to define a 1980 state-of-the-art design of orbital transfer vehicles are summarized. The advanced thrust chamber, turbomachinery, and engine power cycle technologies recommended for development and verification during the 1980's are discussed. The state-of-the-art design serves as one of the reference points for evaluation of the advanced technology concepts. A second reference point is provided by the 1960 state-of-the-art RL-10 engine. Advanced heat energy extraction concepts in the combustor and injector are presented that will permit the high chamber pressures expected of the advanced expander cycle engine concept. Advanced turbomachinery concepts are selected that will help utilize efficiently the heat energy extracted in the thrust chamber and aid in uprating chamber pressures to values five times those realized with 1960 to 1980 technologies. The higher chamber pressure benefits are maximized with nozzle concepts that will package within a short retracted length and extend to the very high expansion area ratios required for specific impulse values significantly higher than those now realized by operational systems. Control system component concepts studied will allow utilization of the high turbine drive energy levels in a reliable repeatable manner that will reduce program risk and enhance overall space transportation system mission life cycle cost.

Glass, J.↗
NASA NTRS1985

Ignition characterization of LOX/hydrocarbon propellants

The results of an evaluation of the ignition characteristics of the gaseous oxygen (Gox)/Ethanol propellant combination are summarized. Ignition characterization was accomplished through the analysis, fabrication and testing of a spark initiated torch igniter and prototype 620 1bF thruster/igniter assembly. The igniter was tested over a chamber pressure range of 3.3 to 262 psia and a mixture ratio range of 0.4 to 40. The prototype thruster/igniter assembly was tested over the chamber pressure range of 74 to 197 psia and mixture ratio range of 0.778 to 3.29. Cold (-92 to -165 F) and ambient (44 to 80 F) propellant temperatures were used. Spark igniter ignition limits and thruster steady state and pulse mode, performance, cooling and stability data are presented. Spark igniter ignition limits are presented in terms of cold flow pressure, ignition chamber diameter and mixture ratio. Thruster performance is presented in terms of vacuum specific impulse versus engine mixture ratio. Gox/Ethanol propellants were shown to be ignitable over a wide range of mixture ratios. Cold propellants were shown to have a minor effect on igniter ignition limits. Thruster pulse mode capability was demonstrated with multiple pulses of 0.08 sec duration and less.

Lawver, B. R.↗
NASA NTRS1989

Transpiration And Regenerative Cooling Of Rocket Engine

Transpiration cooling extends limits of performance. Addition of transpiration cooling to regeneratively-cooled rocket-engine combustion chamber proposed. Modification improves performance of engine by allowing use of higher chamber pressure. Throat section of combustion-chamber wall cooled by transpiration, while chamber and nozzle sections cooled by fluid flowing in closed channels. Concept applicable to advanced, high-performance terrestrial engines or some kinds of industrial combustion chambers. With proper design, cooling scheme makes possible to achieve higher chamber pressure and higher overall performance in smaller engine.

Obrien, Charles J.↗