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At least 451 records · Page 25

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NASA NTRS1983

Control techniques to improve Space Shuttle solid rocket booster separation

The present Space Shuttle's control system does not prevent the Orbiter's main engines from being in gimbal positions that are adverse to solid rocket booster separation. By eliminating the attitude error and attitude rate feedback just prior to solid rocket booster separation, the detrimental effects of the Orbiter's main engines can be reduced. In addition, if angular acceleration feedback is applied, the gimbal torques produced by the Orbiter's engines can reduce the detrimental effects of the aerodynamic torques. This paper develops these control techniques and compares the separation capability of the developed control systems. Currently with the worst case initial conditions and each Shuttle system dispersion aligned in the worst direction (which is more conservative than will be experienced in flight), the solid rocket booster has an interference with the Shuttle's external tank of 30 in. Elimination of the attitude error and attitude rate feedback reduces that interference to 19 in. Substitution of angular acceleration feedback reduces the interference to 6 in. The two latter interferences can be eliminated by atess conservative analysis techniques, that is, by using a root sum square of the system dispersions.

Tomlin, D. D.↗

NASA NTRS1983

Observations of LHR noise with banded structure by the sounding rocket S29 barium-GEOS

The measurement of electrostatic noise near the lower hybrid frequency made by the sounding rocket S29 barium-GEOS is reported. The noise is related to the spin of the rocket and reaches well below the local lower hybrid resonance frequency. Above the altitude of 300 km the noise shows banded structure roughly organized by the hydrogen cyclotron frequency. Simultaneously with the banded structure a signal near the hydrogen cyclotron frequency is detected. This signal is also spin modulated. The character of the noise strongly suggests that it is locally generated by the rocket payload disturbing the plasma. If this interpretation is correct, plasma wave experiments on other spacecrafts are expected to observe similar phenomena.

Koskinen, H. E. J.↗

NASA NTRS1985

The Solid Rocket Booster Auxiliary Power Unit: Meeting the Challenge

The thrust vector control systems of the solid rocket boosters are turbine-powered, electrically controlled hydraulic systems which function through hydraulic actuators to gimbal the nozzles of the solid rocket boosters and provide vehicle steering for the Space Shuttle. Turbine power for the thrust vector control systems is provided through hydrazine fueled auxiliary power units which drive the hydraulic pumps. The solid rocket booster auxiliary power unit resulted from trade studies which indicated significant advantages would result if an existing engine could be found to meet the program goal of 20 missions reusability and adapted to meet the seawater environments associated with ocean landings. During its maturation, the auxiliary power unit underwent many design iterations and provided its flight worthiness through full qualification programs both as a component and as part of the thrust vector control system. More significant, the auxiliary power unit has successfully completed six Shuttle missions.

Hughes, R. W.↗

NASA NTRS1985

Advanced research and technology programs for advanced high-pressure oxygen-hydrogen rocket propulsion

A research and technology program for advanced high pressure, oxygen-hydrogen rocket propulsion technology is presently being pursued by the National Aeronautics and Space Administration (NASA) to establish the basic discipline technologies, develop the analytical tools, and establish the data base necessary for an orderly evolution of the staged combustion reusable rocket engine. The need for the program is based on the premise that the USA will depend on the Shuttle and its derivative versions as its principal Earth-to-orbit transportation system for the next 20 to 30 yr. The program is focused in three principal areas of enhancement: (1) life extension, (2) performance, and (3) operations and diagnosis. Within the technological disciplines the efforts include: rotordynamics, structural dynamics, fluid and gas dynamics, materials fatigue/fracture/life, turbomachinery fluid mechanics, ignition/combustion processes, manufacturing/producibility/nondestructive evaluation methods and materials development/evaluation. An overview of the Advanced High Pressure Oxygen-Hydrogen Rocket Propulsion Technology Program Structure and Working Groups objectives are presented with highlights of several significant achievements.

Marsik, S. J.↗

NASA NTRS1985

Coated oxidizers for combustion stability in solid-propellant rockets

Experiments are conducted in a laboratory-scale (6.25-cm diameter) end-burning rocket motor with state-of-the-art, ammonium perchlorate hydroxy-terminated polybutadiene (HTPB), nonmetallized propellants. The concept of tailoring the stability characteristics with a small amount (less than 1 percent by weight) of COATING on the oxidizer is explored. The thermal degradation characteristics of the coat chemical are deduced through theoretical arguments on thermal diffusivity of the composite material (propellant). Several candidate coats are selected and propellants are cast. These propellants (with coated oxidizers) are fired in a laboratory-scale end-burning rocket motor, and real-time pressure histories are recorded. The control propellant (with no coating) is also tested for comparison. The uniformity of the coating, confirmed by SEM pictures and BET adsorption measurements, is thought to be an advance in technology. The frequency of bulk mode instability (BMI), the pressure fluctuation amplitudes, and stability boundaries are correlated with parameters related to the characteristic length (L-asterisk) of the rocket motor. The coated oxidizer propellants, in general, display greater combustion stability than the control (state-of-the-art). The correlations of the various parameters are thought to be new to a field filled with much uncertainty.

Helmy, A. M.↗

NASA NTRS1985

Advanced research and technology program for advanced high pressure oxygen-hydrogen rocket propulsion

Marsik, S. J.↗

NASA NTRS1985

Launching rockets and small satellites from the lunar surface

Scientific payloads and their propulsion systems optimized for launch from the lunar surface differ considerably from their counterparts for use on earth. For spin-stabilized payloads, the preferred shape is a large diameter-to-length ratio to provide stability during the thrust phase. The rocket motor required for a 50-kg payload to reach an altitude of one lunar radius would have a mass of about 41 kg. To place spin-stabilized vehicles into low altitude circular orbits, they are first launched into an elliptical orbit with altitude about 840 km at aposelene. When the spacecraft crosses the desired circular orbit, small retro-rockets are fired to attain the appropriate direction and speed. Values of the launch angle, velocity increments, and other parameters for circular orbits of several altitudes are tabulated. To boost a 50-kg payload into a 100-km altitude circular orbit requires a total rocket motor mass of about 90 kg.

Anderson, K. A.↗

NASA NTRS1986

Computational simulation of liquid rocket injector anomalies

A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

Przekwas, A. J.↗

NASA NTRS1986

Acceleration effects in solid propellant rocket motors

The performance variations due to acceleration loads imposed on spinning solid propellant rocket motors are investigated. The four potentially most significant modes of acceleration-induced phenomena are identified from a study of the literature and modeled. The four modes are a mechanical mode which deals with deformations of the propellant and case: a thermodynamic mode which covers acceleration-induced combustion phenomena; a stress mode which covers the stressed propellant's effect on burn rate; and a gas dynamic mode which deals with changes in gas flow in the chamber and through the nozzle. Simplified models of each mode are developed or taken from the literature and are added to an internal ballistics evaluation computer program. The resulting analysis is the first to include all of the modes. In order to do this an original analysis of the mechanical and stress modes was necessary. However, the analysis shows that the stress mode is not important for the circular perforated grains studied. The other effects are shown to have a significant influence on solid rocket motor performance. The magnitude of the different mode effects are such that one may not be ignored over the others as has been done in the past. The results of the analysis are compared to published rocket motor data. The comparisons indicate an erosive burning effect that is a function of spin rate. A qualitative explanation of the erosive effect is presented.

Langhenry, M. T.↗

NASA NTRS1987

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m2 (350 psia)

The joint Army. Navy, NASA. Air Force (JANNAF) rocket engine peformnace prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure incorporated into the December l984 version of the TDK program for the high-area-ratio rocket engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the experimental test conditions and TDK input parameters is given. The results show that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values. An experimental value for inviscid thrust was obtained for the nozzle extension between area ratios of 427.5 and 1030 by using an integration of the measured wall static pressures. Subtracting the measured thrust gain produced by the nozzle between area ratios of 427.5 and 1030 from the inviscid thrust gain yielded experimental drag decrements of 10.85 and 27.00 N (2.44 and 6.07 lb) for mixture ratios of 3.04 and 4.29, respectively. These values correspond to 0.45 and 1.11 percent of the total vacuum thrust. At a mixture ratio of 4.29, the TDK predicted drag decrement was 16.59 N (3.73 lb), or 0.71 percent of the predicted total vacuum thrust.

Smith, Tamara A.↗

NASA NTRS1987

Active vehicle charging measurements in sounding rocket and space shuttle orbiter environments at Low Earth Orbit (LEO) altitude

It was concluded that for electron beam emission up to 100mA, vehicle charging is not a significant problem with the Space Shuttle Orbiter. Similarly, sounding rocket payloads have no serious charging problems up to this level of beam current, provided that the maximum amount of the rocket skin is available to collect ionospheric electrons from the LEO altitude range. However, sounding rockets are marginal in their collecting area capability and other effects may occur to balance the beam current when operated at lower altitudes during the night.

Raitt, W. J.↗

NASA NTRS1987

A detailed description of the uncertainty analysis for High Area Ratio Rocket Nozzle tests at the NASA Lewis Research Center

A preliminary uncertainty analysis has been performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis are presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

Kenneth J. Davidian↗

NASA NTRS1986

Results from a series of US/Japan tethered rocket experiments

Tethered rocket experiments have been carried out four times during a US-Japan joint space program in progress since 1980. The goal of the rocket program has been to perform a new type of active experiment by ejecting an electron beam from the tethered mother-daughter payload system. In the third and fourth rocket flights, the conductive tether wire was deployed more than 400 m. It was found that the tether wire acted as an antenna and its antenna impedance decreased with the extension of the wire both in HF and VLF bands. The vehicle charging due to the beam emission up to 80 mA was repeatedly measured in the series of the experiments. During the 80 mA emission, a clear evidence for the ignition of a beam plasma discharge was obtained by the plasma probe, photometers and wave receivers.

Sasaki, S.↗

NASA NTRS1987

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/sq m (350 psia)

The Joint Army, Navy, NASA, Air Force (JANNAF) rocket-engine performance-prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure that has been incorporated into the December 1984 version of the TDK program for the high-area-ratio rocket-engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values.

Smith, Tamara A.↗

NASA NTRS1986

Rocket nozzle thermal shock tests in an arc heater facility

A rocket motor nozzle thermal structural test technique that utilizes arc heated nitrogen to simulate a motor burn was developed. The technique was used to test four heavily instrumented full-scale Star 48 rocket motor 2D carbon/carbon segments at conditions simulating the predicted thermal-structural environment. All four nozzles survived the tests without catastrophic or other structural failures. The test technique demonstrated promise as a low cost, controllable alternative to rocket motor firing. The technique includes the capability of rapid termination in the event of failure, allowing post-test analysis.

Painter, James H.↗

NASA NTRS1987

A detailed description of the uncertainty analysis for high area ratio rocket nozzle tests at the NASA Lewis Research Center

A preliminary uncertainty analysis was performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis is presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

Davidian, Kenneth J.↗

NASA NTRS1987

Reusable rocket engine optical condition monitoring

Plume emission spectrometry and optical leak detection are described as two new applications of optical techniques to reusable rocket engine condition monitoring. Plume spectrometry has been used with laboratory flames and reusable rocket engines to characterize both the nominal combustion spectra and anomalous spectra of contaminants burning in these plumes. Holographic interferometry has been used to identify leaks and quantify leak rates from reusable rocket engine joints and welds.

Wyett, L.↗

NASA NTRS1988

Liquid rocket boosters for the Space Shuttle - An outline of trade issues and analysis techniques

Numerous liquid rocket booster configurations were studied to replace the Solid Rocket Boosters on the current Space Shuttle. While improved safety, reliability, performance and economy of operation were sought, the designs also sought to minimize volumetric growth to reduce aerodynamic loads and structural modifications. Aero loads could be significant as a function of booster diameter if extrapolated directly from solid rocket characteristics. Procedures to meet these goals with analytical booster sizing and simple dynamic simulations are shown. Engine chamber pressure proved to be a good independent variable for implications of pressure-fed and pump-fed engines, fuel and material selection. For some configurations, ascent trajectory modifications were suggested: longer vertical rise, lofting, heads-up orientation, and higher first and second stage target altitudes. Since engine manufacture would dominate booster costs, booster simplification of reuse are discussed based on cost estimates.

Kelly, W. D.↗