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At least 1,405 records · Page 78
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Performance of a low-power subsonic-arc-attachment arcjet thruster

A subsonic-arc-attachment thruster design was scaled from a 30 kW 1960's vintage thruster to operate at nominally 3 kW. Performance measurements were obtained over a 1-4 kW power range using hydrogen as the propellant. Several modes of operation were identified and were characterized by varying degrees of voltage instability. A stability map was developed showing that the voltage oscillations were brought upon by elevated current or propellant levels. At a given specific energy level the specific impulse increased asymptotically with increased flow rates. Comparisons of performance were made between radial and tangential propellant injection. When the vortex flow was eliminated using radial injection, the operating voltages were lower at a given current, and the specific impulse and efficiency decreased. Tests were also conducted to determine the effects of background pressure on operation, and performance data were obtained at pressures of 0.047 Pa and 18 Pa. For a given specific energy level, the performance increased with a decrease in facility background pressure. Lowering the background pressure also caused a dramatic change in the voltage-current characteristic and the voltage stability, a phenomenon not previously reported with conventional supersonic-arc-attachment thrusters.

Sankovic, John M.↗
NASA NTRS1995

PPT Thrust Stand

A torsional-type thrust stand has been designed and built to test Pulsed Plasma Thrusters (PPT's) in both single shot and repetitive operating modes. Using this stand, momentum per pulse was determined strictly as a function of thrust stand deflection, spring constant, and natural frequency. No empirical corrections were required. The accuracy of the method was verified using a swinging impact pendulum. Momentum transfer data between the thrust stand and the pendulum were consistent to within 1%. Following initial calibrations, the stand was used to test a Lincoln Experimental Satellite (LES-8/9) thruster. The LES-8/9 system had a mass of approximately 7.5 kg, with a nominal thrust to weight ratio of 1.3 x 10(exp -5). A total of 34 single shot thruster pulses were individually measured. The average impulse bit per pulse was 266 microN-s, which was slightly less than the value of 300 microN-s published in previous reports on this device. Repetitive pulse measurements were performed similar to ordinary steady-state thrust measurements. The thruster was operated for 30 minutes at a repetition rate of 132 pulses per minute and yielded an average thrust of 573 microN. Using average thrust, the average impulse bit per pulse was estimated to be 260 microN-s, which was in agreement with the single shot data. Zero drift during the repetitive pulse test was found to be approximately 1% of the measured thrust.

Haag, Thomas W.↗
NASA NTRS1995

Performance of a Miniaturized Arcjet

Performance measurements were obtained and life-limiting mechanisms were identified on a laboratory-model arcjet thruster designed to operate at a nominal power level of 300 W. The design employed a supersonic-arc-attachment concept and was operated from 200 to 400 W on hydrogen/nitrogen mixtures in ratios simulating fully decomposed hydrazine and ammonia. Power was provided by breadboard power processor. Performance was found to be a strong function of propellant flow rate. Anode losses were essentially constant for the range of mass flow rates tested. It is believed that the performance is dominated by viscous effects. Significantly improved performance was noted with simulated ammonia operation. At 300 W the specific impulse on simulated ammonia was 410 s with an efficiency of 0.34, while simulated hydrazine provided 370 s specific impulse at an efficiency of 0.27.

Sankovic, John M.↗
NASA NTRS1996

Pulsed Plasma Thrusters for Small Spacecraft Attitude Control

Pulsed plasma thrusters (PPT's) are a new option for attitude control of a small spacecraft and may result in reduced attitude control system (ACS) mass and cost. The primary purpose of an ACS is to orient the spacecraft configuration to the desired accuracy in inertial space. The ACS functions for which the PPT system will be analyzed include disturbance torque compensation and slewing maneuvers such as sun acquisition for which the small impulse bit and high specific impulse of the PPT offers unique advantages. The NASA Lewis Reserach Center (LeRC) currently has a contracted flight PPT system development program in place with Olin Aerospace and a delivery date of October 1997. The PPT system in this study are based upon the work being done under the NASA LeRC program. Analysis of the use of PPT's for ACS showed that the replacement of the standard momentum wheels and torque rods systems with a PTT system to perform the altitude control maneuvers on a small low Earth orbiting spacecraft reduced the ACS mass by 50 to 75 percent with no increase in required power level over comparable wheel-based systems.

McGuire, Melissa L.↗
NASA NTRS1996

Pulsed plasma thrusters for small spacecraft attitude control

Pulsed Plasma Thrusters (PPTS) are a new option for attitude control of a small spacecraft and may result in reduced attitude control system (ACS) mass and cost. The primary purpose of an ACS is to orient the spacecraft to the desired accuracy in inertial space. The ACS functions for which the PPT system will be analyzed include disturbance torque compensation, and slewing maneuvers such as sun acquisition for which the small impulse bit and high specific impulse of the PPT offers unique advantages. The NASA Lewis Research Center (LERC) currently has a contracted flight PPT system development program in place with Olin Aerospace with a delivery date of October 1997. The PPT systems in this study are based upon the work being done under the NASA LERC program. Analysis of the use of PPTs for ACS showed that the replacement of the standard momentum wheels and torque rods with a PPT system to perform the attitude control maneuvers on a small low Earth orbiting spacecraft reduced the ACS mass by 50 to 75% with no increase in required power level over comparable wheel-based systems, though rapid slewing power requirements may present an issue.

McGuire, Melissa L.↗
NASA NTRS1996

Evaluation of Low Power Hall Thruster Propulsion

Hall thruster systems based on the SPT-50 and the TAL D-38 were evaluated and mission studies were performed. The 0.3 kilowatt SPT-50 operated with a specific impulse of 1160 seconds and an efficiency of 0.32. The 0.8 kilowatt D-38 provided a specific impulse above 1700 seconds at an efficiency of 0.5. The D-38 system was shown to offer a 56 kilogram propulsion system mass savings over a 101 kilogram hydrazine monopropellant system designed to perform North-South station keeping maneuvers on board a 430 kilogram geostationary satellite. The SPIT-50 system offered a greater than 50% propulsion system mass reduction in comparison to the chemical system on board a 200 kilogram low Earth orbit spacecraft performing two orbit raises and drag makeup over two years. The performance characteristics of the SPF-50 were experimentally evaluated at a number of operating conditions. The ion current density distribution of this engine was measured. The performance and system mass benefits of advanced systems based on both engines were considered.

Manzella, David↗
NASA NTRS1996

Mission Operations and Data Analysis Phase of the SupraThermal through Energetic Particle detector (STEP) of the EPACT experiment on the Wind Spacecraft

Scientific studies of energetic particle populations observed by STEP are summarized. Science highlights cover: upstream He-3 events general upstream events, energy spectra of ions accelerated in impulsive and gradual solar events, impulsive solar particle events, and corotating interaction regions. Bibliographies of journal publications and papers presented at conferences are also provided.

Mason, G. M.↗
NASA NTRS1997

The duration-photon energy relation in gamma-ray bursts and its interpretations

A correlation between the spectral and temporal structure in gamma ray bursts was presented elsewhere, where it was discovered that the duration of the constituent subpulses of the time profile of a given gamma ray burst have a well-defined power law dependence, of approximately index 0.45, on the energy of the observed photons. Two models are presented which account for the observed correlation. These models involve: the impulsive injection of a population of relativistic electrons; their subsequent cooling by synchrotron radiation; the impulsive injection of mono-energetic high energy photons in a medium of a Thomson depth of approximately 5, and their subsequent downgrading in energy due to electron scattering. Arguments are presented for distinguishing between these two models from the existing data.

Kazanas, Demosthenes↗
NASA NTRS1997

Correlation Filtering of Modal Dynamics using the Laplace Wavelet

Wavelet analysis allows processing of transient response data commonly encountered in vibration health monitoring tasks such as aircraft flutter testing. The Laplace wavelet is formulated as an impulse response of a single mode system to be similar to data features commonly encountered in these health monitoring tasks. A correlation filtering approach is introduced using the Laplace wavelet to decompose a signal into impulse responses of single mode subsystems. Applications using responses from flutter testing of aeroelastic systems demonstrate modal parameters and stability estimates can be estimated by correlation filtering free decay data with a set of Laplace wavelets.

Freudinger, Lawrence C.↗
NASA NTRS1998

Scaling of Ion Thrusters to Low Power

Analyses were conducted to examine ion thruster scaling relationships in detail to determine performance limits, and lifetime expectations for thruster input power levels below 0.5 kW. This was motivated by mission analyses indicating the potential advantages of high performance, high specific impulse systems for small spacecraft. The design and development status of a 0.1-0.3 kW prototype small thruster and its components are discussed. Performance goals include thruster efficiencies on the order of 40% to 54% over a specific impulse range of 2000 to 3000 seconds, with a lifetime in excess of 8000 hours at full power. Thruster technologies required to achieve the performance and lifetime targets are identified.

Patterson, Michael J.↗
NASA NTRS1960

Experimental Performance of Area Ratio 200, 25 and 8 Nozzles on JP-4 Fuel and Liquid Oxygen Rocket Engine

The performance of an area ratio 200 bell-shaped nozzle, an area ratio 25 bell-shaped nozzle, and an area ratio 8 conic nozzle on a JP-4 fuel and liquid-oxygen rocket engine has been determined. Tests were conducted using a nominal 4000-pound-thrust rocket in the Lewis 10- by 10-foot supersonic tunnel, which provided the altitude environment needed for fully expanded nozzle flow. The area ratio 200 nozzle had a vacuum thrust coefficient of 1.96, compared with 1.82 and 1.70 for the area ratio 25 and 8 nozzles, respectively. These values are approximately equal to those for theoretical frozen expansion. The measured value of vacuum specific impulse for the area ratio 200 nozzle was 317 seconds for a combustion-chamber characteristic velocity of 5200 feet per second. The vacuum-specific-impulse increase for the area-ratio increase from 8 to 200 was 46 seconds.

Lovell, J. Calvin↗
NASA NTRS1959

A General Method for Automatic Computation of Equilibrium Compositions and Theoretical Rocket Performance of Propellants

A general computer program for chemical equilibrium and rocket performance calculations was written for the IBM 650 computer with 2000 words of drum storage, 60 words of high-speed core storage, indexing registers, and floating point attachments. The program is capable of carrying out combustion and isentropic expansion calculations on a chemical system that may include as many as 10 different chemical elements, 30 reaction products, and 25 pressure ratios. In addition to the equilibrium composition, temperature, and pressure, the program calculates specific impulse, specific impulse in vacuum, characteristic velocity, thrust coefficient, area ratio, molecular weight, Mach number, specific heat, isentropic exponent, enthalpy, entropy, and several thermodynamic first derivatives.

Gordon, Sanford↗
NASA NTRS1961

Theoretical Performance of Hydrogen-Oxygen Rocket Thrust Chambers

Data are presented for liquid-hydrogen-liquid-oxygen thrust chambers at chamber pressures from 15 to 1200 pounds per square inch absolute, area ratios to approximately 300, and percent fuel from about 8 to 34 for both equilibrium and frozen composition during expansion. Specific impulse in vacuum, specific impulse, combustion-chamber temperature, nozzle-exit temperature, characteristic velocity, and the ratio of chamber-to-nozzle-exit pressure are included. The data are presented in convenient graphical forms to allow quick calculation of theoretical nozzle performance with over- or underexpansion, flow separation, and introduction of the propellants at various initial conditions or heat loss from the combustion chamber.

Sievers, Gilbert K.↗
NASA NTRS1960

Flight Performance of a Spin-Stabilized 20-Inch-Diameter Solid-Propellant Spherical Rocket Motor

A successful flight test of a spin-stabilized 20-inch-diameter solid-propellant rocket motor having a propellant mass fraction of 0.92 has been made. The motor was fired at altitude after being boosted by a three-stage test vehicle. Analysis of the data indicates that a total impulse of 44,243 pound-second with a propellant specific impulse of approximately 185 was achieved over a total action time of about 12 seconds. These results are shown to be in excellent agreement with data from ground static firing tests of these motors. The spherical rocket motor with an 11-pound payload attained a velocity of 15,620 feet per second (m = 16.7) with an incremental velocity increase for the spherical motor stage of 12,120 feet per second.

Levine, Jack↗
NASA NTRS1959

Theoretical Performance of Liquid Hydrogen with Liquid Oxygen as a Rocket Propellant

Theoretical rocket performance for both equilibrium and frozen composition during expansion was calculated for the propellant combination liquid hydrogen and liquid oxygen at four chamber pressures (60, 150, 300, and 600 lb/sq in. abs) and a wide range of pressure ratios (1 to 4000) and oxidant-fuel ratios (1.190 to 39.683). Data are given to estimate performance parameters at chamber pressures other than those for which data are tabulated. The parameters included are specific impulse, specific impulse in vacuum, combustion-chamber temperature, nozzle-exit temperature, molecular weight, molecular-weight derivatives, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, Mach number, and equilibrium gas compositions.

Gordon, Sanford↗
NASA NTRS1998

Input Shaping to Reduce Solar Array Structural Vibrations

Structural vibrations induced by actuators can be minimized using input shaping. Input shaping is a feedforward method in which actuator commands are convolved with shaping functions to yield a shaped set of commands. These commands are designed to perform the maneuver while minimizing the residual structural vibration. In this report, input shaping is extended to stepper motor actuators. As a demonstration, an input-shaping technique based on pole-zero cancellation was used to modify the Solar Array Drive Assembly (SADA) actuator commands for the Lewis satellite. A series of impulses were calculated as the ideal SADA output for vibration control. These impulses were then discretized for use by the SADA stepper motor actuator and simulated actuator outputs were used to calculate the structural response. The effectiveness of input shaping is limited by the accuracy of the knowledge of the modal frequencies. Assuming perfect knowledge resulted in significant vibration reduction. Errors of 10% in the modal frequencies caused notably higher levels of vibration. Controller robustness was improved by incorporating additional zeros in the shaping function. The additional zeros did not require increased performance from the actuator. Despite the identification errors, the resulting feedforward controller reduced residual vibrations to the level of the exactly modeled input shaper and well below the baseline cases. These results could be easily applied to many other vibration-sensitive applications involving stepper motor actuators.

Doherty, Michael J.↗
NASA NTRS1963

Performance Capability of Single-Cavity Vortex Gaseous Nuclear Rockets

An analysis was made to determine the maximum powerplant thrust-to-weight ratio possible with a single-cavity vortex gaseous reactor in which all the hydrogen propellant must diffuse through a fuel-rich region. An assumed radial temperature profile was used to represent conduction, convection, and radiation heat-transfer effects. The effect of hydrogen property changes due to dissociation and ionization was taken into account in a hydrodynamic computer program. It is shown that, even for extremely optimistic assumptions of reactor criticality and operating conditions, such a system is limited to reactor thrust-to-weight ratios of about 1.2 x 10(exp -3) for laminar flow. For turbulent flow, the maximum thrust-to-weight ratio is less than 10(exp -3). These low thrusts result from the fact that the hydrogen flow rate is limited by the diffusion process. The performance of a gas-core system with a specific impulse of 3000 seconds and a powerplant thrust-to-weight ratio of 10(exp -2) is shown to be equivalent to that of a 1000-second advanced solid-core system. It is therefore concluded that a single-cavity vortex gaseous reactor in which all the hydrogen must diffuse through the nuclear fuel is a low-thrust device and offers no improvement over a solid-core nuclear-rocket engine. To achieve higher thrust, additional hydrogen flow must be introduced in such a manner that it will by-pass the nuclear fuel. Obviously, such flow must be heated by thermal radiation. An illustrative model of a single-cavity vortex system employing supplementary flow of hydrogen through the core region is briefly examined. Such a system appears capable of thrust-to-weight ratios of approximately 1 to 10. For a high-impulse engine, this capability would be a considerable improvement over solid-core performance. Limits imposed by thermal radiation heat transfer to cavity walls are acknowledged but not evaluated. Alternate vortex concepts that employ many parallel vortices to achieve higher hydrogen flow rates offer the possibility of sufficiently high thrust-to-weight ratios, if they are not limited by short thermal-radiation path lengths.

Ragsdale, Robert G.↗
NASA NTRS1998

Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

Smith, Timothy D.↗