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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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Saturn I /SA-8/ launch vehicle and launch complex systems - H- 1 rocket engine, and hydraulic systems
NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Test methodology and conditions are presented, and base heating results from 76 runs are reported in non-dimensional form. Regions of high heating are identified and comparisons of various configuration and conditions are highlighted. Base pressure and radiometer results are also reported.
IBM 360/40 computer programs on preigniter rocket engine hydraulics and steady state performance of Apollo SM-LM RCS engine
The Space Shuttle Orbit Maneuvering Subsystem Rocket Engine employs a platelet element injector concept. This injector has demonstrated 316-sec vacuum specific impulse performance under simulated altitude conditions when tested with a milled slot/electroformed nickel close-out regenerative chamber and a full 71 area ratio nozzle. To date, over 300 altitude engine tests and 300 stability bomb tests have demonstrated stable, erosion free operation with this concept to test durations of 150 seconds. The injector and chamber also meet the reusable requirements of the shuttle with a cycle life capability in excess of 1000 cycles. An extensive altitude restart program has also demonstrated OMS-engine operation over large variations in the burn and coast times with helium saturated propellants.
J-2X engine testing on the A-2 test stand at the NASA John C. Stennis Space Center (SSC) has recently concluded. As part of that test campaign, the engine was operated at lower power levels in support of expanding the use of J-2X to other missions. However, the A-2 diffuser was not designed for engine testing at the proposed low power levels. To evaluate the risk of damage to the diffuser, computer simulations were created of the rocket engine exhaust plume inside the 50ft long, water-cooled, altitude-simulating diffuser. The simulations predicted that low power level testing would cause the plume to oscillate in the lower sections of the diffuser. This can possibly cause excessive vibrations, stress, and heat transfer from the plume to the diffuser walls. To understand and assess the performance of the diffuser during low power level engine testing, nine accelerometers and four strain gages were installed around the outer surface of the diffuser. The added instrumentation also allowed for the verification of the rocket exhaust plume computational model. Prior to engine hot-fire testing, a diffuser water-flow test was conducted to verify the proper operation of the newly installed instrumentation. Subsequently, two J-2X engine hot-fire tests were completed. Hot-Fire Test 1 was 11.5 seconds in duration, and accelerometer and strain data verified that the rocket engine plume oscillated in the lower sections of the diffuser. The accelerometers showed very different results dependent upon location. The diffuser consists of four sections, with Section 1 being closest to the engine nozzle and Section 4 being farthest from the engine nozzle. Section 1 accelerometers showed increased amplitudes at startup and shutdown, but low amplitudes while the diffuser was started. Section 3 accelerometers showed the opposite results with near zero G amplitudes prior to and after diffuser start and peak amplitudes to +/- 100G while the diffuser was started. Hot-Fire Test 1 strain gages showed different data dependent on section. Section 1 strains were small, and were in the range of 50 to 150 microstrain, which would result in stresses from 1.45 to 4.35 ksi. The yield stress of the material, A-285 Grade C Steel, is 29.7 ksi. Section 4 strain gages showed much higher values with strains peaking at 1600 microstrain. This strain corresponds to a stress of 46.41 ksi, which is in excess of the yield stress, but below the ultimate stress of 55 to 75 ksi. The decreased accelerations and strain in Section 1, and the increased accelerations and strain in Sections 3 and 4 verified the computer simulation prediction of increased plume oscillations in the lower sections of the diffuser. Hot-Fire Test 2 ran for a duration of 125 seconds. The engine operated at a slightly higher power level than Hot-Fire Test 1 for the initial 35 seconds of the test. After 35 seconds the power level was lowered to Hot-Fire Test 1 levels. The acceleration and strain data for Hot-Fire Test 2 was similar during the initial part of the test. However, just prior to the engine being lowered to the Hot-Fire Test 1 power level, the strain gage data in Section 4 showed a large decrease to strains near zero microstrain from their peak at 1500 microstrain. Future work includes further strain and acceleration data analysis and evaluation.
NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.
A conceptual level analysis was performed on a horizontal-takeoff, two-stage-to-orbit system consisting of a rocket and turboramjet powered hypersonic booster and a rocket powered orbiter. The analysis includes estimates of vehicle aerodynamics, performance, and weights. Rocket and airbreathing systems on the booster are operated in parallel and at full thrust for ascent to a Mach 6 staging point, while cruise-back is accomplished using airbreathing propulsion alone. Rocket engines on the orbiter are ignited at staging to propel the orbiter. Booster weights were determined for various combinations of rocket and airbreathing propulsion. The combinations that resulted in the lowest vehicle gross and empty weights were determined. Results show that the lowest gross weight occurs when only airbreathing engines are used. The empty weight of the all airbreathing booster, however, is quite high. Significant reductions in booster empty weight is accomplished with the addition of rocket engines. Also, the use of rockets on the booster greatly decreases the time to staging and the cruise-back distance.
Economical feasibility of open and closed cycle vortex-stabilized gaseous nuclear rocket engines, investigating vortex flow characteristics
Chlorine and oxides of nitrogen (NO(x)) released into the atmosphere contribute to acid rain (ground level or low-altitude sources) and ozone depletion from the stratosphere (high-altitude sources). Rocket engines have the potential for forming or activating these pollutants in the rocket plume. For instance, H2/O2 rockets can produce thermal NO(x) in their plumes. Emphasis, in the past, has been placed on determining the impact of chlorine release on the stratosphere. To date, very little, if any, information is available to understand what contribution NO(x) emissions from ground-based engine testing and actual rocket launches have on the atmosphere. The goal of this work is to estimate the afterburning emissions from chemical rocket plumes and determine their local stratospheric impact. Our study focuses on the space shuttle rocket motors, which include both the solid rocket boosters (SRB's) and the liquid propellant main engines (SSME's). Rocket plume afterburning is modeled employing a one-dimensional model incorporating two chemical kinetic systems: chemical and thermal equilibria with overlayed nitric oxide chemical kinetics (semi equilibrium) and full finite-rate chemical kinetics. Additionally, the local atmospheric impact immediately following a launch is modeled as the emissions diffuse and chemically react in the stratosphere.
Radiative heat transfer to the propellant and reactor criticality for a fissionable gaseous rocket engine are analyzed to determine their interdependence. The necessity for propellant thicknesses of approximately 1-3 m due to poor thermal absorption properties of hydrogen significantly affects reactor critical radius and mass. The two primary adverse effects are: (1) increased absorption in the reflector- moderator for a given reflector thickness and (2) poor utilization of thermal neutrons by the core due to the lower geometrical view factor of the core for the reflector walls. In fact, there is a minimum core radius at a particular propellant thickness which allows the system to “go” critical. Engine performance is limited primarily to two regions of operation: the first, a specific impulse of approximately 1550 sec at a thrust level of 2 x 10(exp 6) lb. and second, a specific impulse of approximately 2200 sec at a thrust level of 5.3 x 10(exp 6) lb.
Calculation of spectral transmittance for a particular inhomogeneous gas path is made by combining known data on gases at constant temperature, pressure, and concentration. The spectral transmittances of the inhomogeneous plume gases is needed to calculate the heat radiated from the exhaust plume to the rocket base of a multiple engine rocket.
The NASA Lewis Research Center, located in Cleveland, Ohio has a number of space chemical propulsion test facilities which constitute a significant national space testing resource. The purpose of this paper is to make more users aware of these test facilities and to encourage their use through cooperative agreements between the government, industry, and universities. Research which is of interest to the government is especiallly encouraged and often can be done in a cooperative manner that best uses the resources of all parties. This paper presents an overview of the Lewis test facilities. These facilities are clustered into three test areas: the Rocket Engine Test Facilities (RETF), the Rocket Laboratory (RL), and the Cryogenic Components Laboratory (CCL).
The Real Time Vibration Monitoring System (RTVMS) is a 32-channel high speed vibration data acquisition and processing system developed at Marshall Space Flight Center (MSFC). It Delivers sample rates as high as 51,200 samples/second per channel and performs Fast Fourier Transform (FFT) processing via on-board digital signal processing (DSP) chips in a real-time format. Advanced engine health assessment is achieved by utilizing the vibration spectra to provide accurate sensor validation and enhanced engine vibration redlines. Discrete spectral signatures (such as synchronous) that are indicators of imminent failure can be assessed and utilized to mitigate catastrophic engine failures- a first in rocket engine health assessment. This paper is presented in viewgraph form.
ABCC engine combines air breathing and rocket engine into a single engine to increase the specific impulse over an entire flight trajectory. Except for the heat source, the basic operation of the ABCC is similar to the basic operation of the RBCC engine. The ABCC is intended to have a higher specific impulse than the RBCC for single stage Earth to orbit vehicle. Computational fluid dynamics (CFD) is a useful tool for the analysis of complex transport processes in various components in ABCC propulsion system. The objective of the present research was to develop a transient 1-D numerical model using conservation of mass, linear momentum, and energy equations that could be used to predict flow behavior throughout a generic ABCC engine following a flight path. At specific points during the development of the 1-D numerical model a myriad of tests were performed to prove the program produced consistent, realistic numbers that follow compressible flow theory for various inlet conditions.
A comprehensive mathematical model of mass diffusion has been developed for binary fluids at high pressures, including critical and supercritical pressures. Heretofore, diverse expressions, valid for limited parameter ranges, have been used to correlate high-pressure binary mass-diffusion-coefficient data. This model will likely be especially useful in the computational simulation and analysis of combustion phenomena in diesel engines, gas turbines, and liquid rocket engines, wherein mass diffusion at high pressure plays a major role.
Development is underway of the J -2X engine, a liquid oxygen/liquid hydrogen rocket engine for use on the Space Launch System. The Engine E10001 began hot fire testing in June 2011 and testing will continue with subsequent engines. The J -2X engine main combustion chamber contains both acoustic cavities and baffles. These stability aids are intended to dampen the acoustics in the main combustion chamber. Verification of the engine thrust chamber stability is determined primarily by examining experimental data using a dynamic stability rating technique; however, additional requirements were included to guard against any spontaneous instability or rough combustion. Startup and shutdown chug oscillations are also characterized for this engine. This paper details the stability requirements and verification including low and high frequency dynamics, a discussion on sensor selection and sensor port dynamics, and the process developed to assess combustion stability. A status on the stability results is also provided and discussed.
Scaling in unstable combustion processes of solid propellant rocket engines
Vortex amplifier for controlling secondary flow for thrust vector control of liquid rocket engine