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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Liquid Oxygen Cooling of Hydrocarbon Fueled Rocket Thrust Chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the collant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1989

Liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

Rocket engines using liquid oxygen (LOX) and hydrocarbon fuel as the propellants are being given serious consideration for future launch vehicle propulsion. Normally, the fuel is used to regeneratively cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability. Another possibility for the coolant is the liquid oxygen. Combustion chambers previously tested with LOX and RP-1 as propellants and LOX as the coolant demonstrated the feasibility of using liquid oxygen as a coolant up to a chamber pressure of 13.8 MPa (2000 psia). However, there was concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot gas side wall. In order to study this effect, chambers were fabricated with slots machined upstream of the throat between the cooling passage wall and the hot gas side wall to simulate cracks. The chambers were tested at a nominal chamber pressure of 8.6 MPa (1247 psia) over a range of mixture ratios from 1.9 to 3.1 using liquid oxygen as the coolant. The results of the testing showed that the leaking LOX did not have a deleterious effect on the chambers in the region of the slots. However, there was unexplained melting in the throat region of both chambers, but not in line with the slots.

Armstrong, Elizabeth S.↗
NASA NTRS1999

Mixing Characteristics of Coaxial Injectors at High Gas to Liquid Momentum Ratios

A study of the spray of a swirl coaxial gas-liquid injector operating at high gas to liquid momentum ratios is reported. Mixing and droplet size characteristics of the swirl injector are also compared to a shear coaxial injector, currently being used in the Space Shuttle Main Engine fuel preburner. The injectors were tested at elevated chamber pressures using water as a LOX simulant and nitrogen and helium as gaseous hydrogen simulants. The elevated chamber pressure allowed for matching of several of the preburner injector conditions including; gas to liquid momentum ratio, density ratio and Mach number. Diagnostic techniques used to characterize the spray included; strobe back-light imaging, laser sheet spray imaging, mechanical patternation, and a phase Doppler interferometry. Results thus far indicate that the radial spreading of the swirl coaxial spray is much less than was reported in previous studies of swirl injectors operating at atmospheric back-pressure. The swirl coaxial spray does, however, exhibit a smaller overall droplet size which may be interpreted as an increase in local mixing.

Strakey, P. A.↗
NASA NTRS1980

High-density fuel combustion and cooling investigation

The analysis, design, fabrication and testing of several engine configurations are discussed with respect to the combustion and heat transfer characteristics of LOX/RP-1 at chamber pressures between 6895 and 13790 kPa (1000 and 2000 psia). The different engine configurations discussed include: 8274 kPa and 13790 kPa (1200 psia and 2000 psia) chamber pressure injectors with like doublet and preatomized triplet elements; cooled and uncooled acoustic resonators; and graphite, regeneratively cooled and calorimetric chambers ranging in length from 27.9 to 37.5 cm (11 to 15 in.). A high pressure LOX/RP-1 spark igniter is also evaluated.

Labotz, R. J.↗
NASA NTRS1993

Performance assessment of low pressure nuclear thermal propulsion

An increase in Isp for nuclear thermal propulsion systems is desirable for reducing the propellant requirements and cost of future applications, such as the Mars Transfer Vehicle. Several previous design studies have suggested that the Isp could be increased substantially with hydrogen dissociation/recombination. Hydrogen molecules (H2), at high temperatures and low pressures, will dissociate to monatomic hydrogen (H). The reverse process (i.e., formation of H2 from H) is exothermic. The exothermic energy in a nozzle increases the kinetic energy and therefore, increases the Isp. The low pressure nuclear thermal propulsion system (LPNTP) system is expected to maximize the hydrogen dissociation/recombination and Isp by operating at high chamber temperatures and low chamber pressures. The process involves hydrogen flow through a high temperature, low pressure fission reactor, and out a nozzle. The high temperature (approximately 3000 K) of the hydrogen in the reactor is limited by the temperature limits of the reactor material. The minimum chamber pressure is about 1 atm because lower pressures decrease the engines thrust to weight ratio below acceptable limits. This study assumes that hydrogen leaves the reactor and enters the nozzle at the 3000 K equilibrium dissociation level. Hydrogen dissociation in the reactor does not affect LPNTP performance like dissociation in traditional chemical propulsion systems, because energy from the reactor resupplies energy lost due to hydrogen dissociation. Recombination takes place in the nozzle due primarily to a drop in temperature as the Mach number increases. However, as the Mach number increases beyond the nozzle throat, the static pressure and density of the flow decreases and minimizes the recombination. The ideal LPNTP Isp at 3000 K and 10 psia is 1160 seconds due to the added energy from fast recombination rates. The actual Isp depends on the finite kinetic reaction rates which affect the amount of monatomic hydrogen recombination before the flow exits the nozzle. A LPNTP system has other technical issues (e.g. flow instability and two-phase flow) besides hydrogen dissociation/recombination which affect the systems practicality. In this study, only the effects of hydrogen dissociation/recombination are examined.

Gerrish, Harrold P., Jr.↗
NASA NTRS1993

Performance assessment of low pressure nuclear thermal propulsion

A low pressure nuclear thermal propulsion (LPNTP) system, which takes advantage of hydrogen dissociation/recombination, was proposed as a means of increasing engine specific impulse (Isp). The effect of hydrogen dissociation/recombination on LPNTP Isp is examined. A two-dimensional computer model was used to show that the optimum chamber pressure is approximately 100 psia (at a chamber temperature of 3,000 K), with an Isp approximately 15 s higher than at 1,000 psia. At high chamber temperatures and low chamber pressures, the increase in Isp is due to both lower average molecular weights caused by dissociation and added kinetic energy from monatomic hydrogen recombination. Monatomic hydrogen recombination increases the Isp more then hydrogen dissociation. Variations in the mole fraction of monatomic hydrogen are similar to variations in static pressure along the axial nozzle position. Most recombination occurs close to the nozzle throat. Practical variations in nozzle geometry have minimal impact on recombination. Other models which can simulate a wider range of nozzle designs should be used in the future. The uncertainty of the hydrogen kinetic reaction rates at high temperatures (approximately 3,000 K) affects the accuracy of the analysis and should be verified with simple bench tests.

Gerrish, H. P., Jr.↗
NASA NTRS2004

Performance and Stability Characteristics of a Uni-Element Swirl Injector for Oxygen-Rich Stage Combustion Cycles

A uni-element liquid propellant combustion performance and instability study for liquid RP-1 and hot oxygen-rich pre-burner products was conducted, at a chamber pressure of about 1000 psi. using flush and recessed swirl injectors. High-frequency pressure transducer measurements were analyzed to yield the characteristic frequencies which were compared to expected frequencies of the chamber. Modes, which were discovered to be present within the main chamber included, the first longitudinal, detected at approximately 1950 Hz, and the second longitudinal mode at approximately 3800 Hz. An additional first longitudinal quarter wave mode was measured at a frequency of approximately 23000 Hz for the recessed swirl injector configuration. The characteristic instabilities resulting from these experiments were relatively weak averaging 0.2% to 0.3% of the chamber pressure.

Pal, S.↗
NASA NTRS2014

Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development & Performance Analysis

ATA-002 Technical Team has successfully designed, developed, tested and assessed the SLS Pathfinder propulsion systems for the Main Base Heating Test Program. Major Outcomes of the Pathfinder Test Program: Reach 90% of full-scale chamber pressure Achieved all engine/motor design parameter requirements Reach steady plume flow behavior in less than 35 msec Steady chamber pressure for 60 to 100 msec during engine/motor operation Similar model engine/motor performance to full-scale SLS system Mitigated nozzle throat and combustor thermal erosion Test data shows good agreement with numerical prediction codes Next phase of the ATA-002 Test Program Design & development of the SLS OML for the Main Base Heating Test Tweak BSRM design to optimize performance Tweak CS-REM design to increase robustness MSFC Aerosciences and CUBRC have the capability to develop sub-scale propulsion systems to meet desired performance requirements for short-duration testing.

Mehta, Manish↗
NASA NTRS1980

Analytical investigation of two hydrogen-oxygen rocket engine systems for low-thrust application

Two hydrogen-oxygen rocket engine system concepts were analyzed parametrically over a thrust range from 100 to 1000 pounds and a chamber pressure range from 175 to 1000 psia. Both concepts were regeneratively cooled with hydrogen and were pump-fed by electric motor driven positive displacement pumps. Electric power was provided by either a turboalternator (turboalternator concept) or some means external to the engine system (auxiliary power concept). The computer program used to conduct the analyses along with the design characteristics of the major engine system components are briefly described. The feasible design range of the systems over the parametric range of thrust is discussed in terms of allowable chamber pressure considering the constraints of thrust chamber cooling and cycle power. Engine system estimated performance, mass, and dimensional envelope parametric data within the feasible design range are presented.

Scheer, D. D.↗
NASA NTRS1974

Fundamental and applied research on core engine/combustion noise of aircraft engines

Some results of a study of the importance of geometrical features of the combustor to combustion roughness and resulting noise are presented. Comparison is made among a perforated can flame holder, a plane slotted flame holder and a plane slotted flame holder which introduces two counter swirling streams. The latter is found to permit the most stable, quiet combustion. Crosscorrelations between the time derivative of chamber pressure fluctuations and far field noise are found to be stronger than between the far field noise and the direct chamber pressure signal. Temperature fluctuations in the combustor nozzle are also found to have a reasonably strong crosscorrelation with far field sound.

Plett, E. G.↗
NASA NTRS1957

Altitude Starting Tests of a 1000-Pound-Thrust Solid-Propellant Rocket

Four solid-propellant rocket engines of nominal 1000-pound-thrust were tested for starting characteristics at pressure altitudes ranging from 112,500 to 123,000 feet and at a temperature of -75 F. All engines ignited and operated successfully. Average chamber pressures ranged from 1060 to ll90 pounds per square inch absolute with action times from 1.51 to 1.64 seconds and ignition delays from 0.070 t o approximately 0.088 second. The chamber pressures and action times were near the specifications, but the ignition delay was almost twice the specified value of 0.040 second.

Sloop, John L.↗
NASA NTRS1993

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

Niiya, Karen E.↗
NASA NTRS1972

Rocket nozzle automatic release system

The function of the mechanism described is to automatically jettison a rocket nozzle element at the conclusion of boost in order to achieve a cruise mode nozzle configuration for the balance of powered flight. The boost nozzle is retained by latched pistons that are mechanically secured until released by chamber pressure, which simultaneously assumes retention of the nozzle by holding the primary piston in place. Upon attainment of cruise mode, the chamber pressure drops, allowing springs to overcome the primary piston force and pull the keeper ring forward to free the nozzle for jettison by motor exhaust.

Kimball, J. B.↗
NASA NTRS1980

Low-thrust chemical rocket engine study

Parametric data and preliminary designs on liquid rocket engines for low thrust cargo orbit-transfer-vehicles are described and those items where technology is required to enhance the designs are identified. The results of film cooling studies to establish the upper chamber pressure limit are given. The study showed that regen cooling with RP-1 was not feasible over the entire thrust and chamber pressure ranges. The thermal data showed that the RP-1 bulk temperature exceeded the study coking temperature limit of 1010 R. Based upon the results presented, O2/H2 and O2/CH4 regen engine systems and O2/H2 film cooled engines were selected for further study in the system analysis. Six engine design concepts are examined.

Mellish, J. A.↗
NASA NTRS1956

Theoretical performance of JP-4 fuel and liquid oxygen as a rocket propellant II : equilibrium composition

Data were calculated for equivalence ratios of 1 to 3, chamber pressures of 300 and 600 pounds per square inch absolute, and pressure ratios of 1 to 1500. Parameters included are specific impulse, combustion and exit temperature, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. A correlation is given which permits determination of performance for a wide range of chamber pressures.

Huff, Vearl N↗
NASA NTRS1956

Theoretical performance of JP-4 fuel with a 70-30 mixture of fluorine and oxygen as a rocket propellant : equilibrium composition

Data were calculated for equivalence ratios of 1 to 4, chamber pressures of 300 and 600 pounds per square inch absolute, and pressure ratios of 1 to 1500. Parameters included are specific impulse, combustion and exit temperatures, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. A correlation is given which permits determination of performance for a wide range of chamber pressures. A method for obtaining specific impulse of JP-4 fuel with OF2 and O3-F2 mixtures is given.

Gordon, Sanford↗
NASA NTRS1973

Experimental investigation of combustor effects on rocket thrust chamber performance

The results are reported of a program to develop special instrumentation systems and engine hardware, conduct tests using LOX/GH2 propellants wherein radial mixtures ratio stratification was controlled, and subsequently compare the results of four selected tests with the predictions of the JANNAF performance-prediction computer programs. During the experiments, the overall propellant mixture ratio was varied from 4.4 to 6.6, while the mixture ratios in the core and outer zone were varied from 5.7 to 8.8 and from 3.7 to 7.2, respectively. A nominal 10 percent of the total fuel flow was used as boundary layer collant in a majority of the firings. Nominal chamber pressure was either 225 or 250 psia, with nozzle expansion ratios of either 25:1 or 4:1. Measurements of the axial chamber pressure and wall heat flux profiles, together with samples of the exhaust gas, were obtained. The corrected experimental specific impulse and characteristic exhaust velocity efficiencies were approximately 97.5 and 98.5 percent, respectively.

Nurick, W. H.↗
NASA NTRS1999

Simulation of Non-Acoustic Combustion Instability in a Hybrid Rocket Motor

A transient model of a hybrid motor was formulated to study the cause and elimination of non-acoustic combustion instability. The transient model was used to simulate four key tests out of a series of seventeen hybrid motor tests conducted by Thiokol, Rocketdyne and Martin Marietta at NASA/Marshall Space Flight Center (NASA/MSFC). These tests were performed under the Hybrid Propulsion Technology for Launch Vehicle Boosters (HPTLVB) program. The first test resulted in stable combustion. The second test resulted in large-amplitude, 6.5 Hz chamber pressure oscillations that gradually damped away by the end of the test. The third test resulted in large-amplitude, 7.5 Hz chamber pressure oscillations that were sustained throughout the test. The seventh test resulted in the elimination of combustion instability with the installation of an orifice immediately upstream of the injector. The formulation and implementation of the model are the scope of this presentation. The current model is an independent continuation of modeling presented previously by joint Thiokol-Rocketdyne collaborators Boardman, Hawkins, Wassom, and Claflin. The previous model simulated an unstable IR&D hybrid motor test performed by Thiokol. There was very good agreement between the model and the test data. Like the previous model, the current model was developed using Matrix-x simulation software. However, the tests performed at NASA/MSFC under the HPTLVB program were actually simulated. In the current model, the hybrid motor consisting of the liquid oxygen (LOX) injector, the multi-port solid fuel grain and the nozzle was simulated. Also, simulated in the model was the LOX feed system consisting of the tank, venturi, valve and feed lines. All components of the hybrid motor and LOX feed system are treated by a lumped-parameter approach. Agreement between the results of the transient model and the actual test data was very good. This agreement between simulated and actual test data indicated that the combustion instability in the hybrid motor was due to two causes. The first cause was a LOX feed system of insufficient stiffness. The second cause was a LOX injector with an impedance or pressure drop that was too low to provide damping against the feed system oscillations. Also, it was discovered that testing with a new grain of solid fuel sustained the combustion instability. However, testing with a used grain of solid fuel caused the combustion instability to gradually decay.

Rocker, Marvin↗