Results for “THROTTLING” — Page 36

NASA NTRS2008

The Development of a Control System for a 5 Kilowatt Free Piston Stirling Engine Convertor

The new NASA Vision for Exploration, announced by President Bush in January 2004, proposes an ambitious program that plans to return astronauts to the moon by the 2018 time frame. A recent NASA study entitled "Affordable Fission Surface Power Study" recommended a 40 kWe, 900 K, NaK-cooled, Stirling convertors for 2020 launch. Use of two of the nominal 5 kW convertors allows the system to be dynamically balanced. A group of four dual-convertor combinations that would yield 40 kWe can be tested to validate the viability of Stirling technology for space fission surface power systems. The work described in this paper deals specifically with the control system for the 5 kW convertor described in the preceding paragraph. This control system is responsible for maintaining piston stroke to a setpoint in the presence of various disturbances including electrical load variations. Pulse starting of the Free Piston Stirling Engine (FPSE) convertor is also an inherent part of such a control system. Finally, the ability to throttle the engine to match the required output power is discussed in terms of setpoint control. Several novel ideas have been incorporated into the piston stroke control strategy that will engender a stable response to disturbances in the presence of midpoint drift while providing useful data regarding the position of both the power piston and displacer.

Kirby, Raymond L.↗

NASA NTRS2008

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Unsteady Analyses and Risk Assessment - Part 2

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Design Support of the feasibility of operating conditions and procedures is critical in such cases due to the possibility of startup/shutdown transients, moving shock structures, unsteady shock-boundary layer interactions and engine and diffuser unstart modes that can result in catastrophic failure. Analyses of such systems is difficult due to resolution requirements needed to accurately capture moving shock structures, shock-boundary layer interactions, two-phase flow regimes and engine unstart modes. In a companion paper, we will demonstrate with the use of CFD, steady analyses advanced capability to evaluate supersonic diffuser and steam ejector performance in the sub-scale A-3 facility. In this paper we will address transient issues with the operation of the facility especially at startup and shutdown, and assess risks related to afterburning due to the interaction of a fuel rich plume with oxygen that is a by-product of the steam ejectors. The primary areas that will be addressed in this paper are: (1) analyses of unstart modes due to flow transients especially during startup/ignition, (2) engine safety during the shutdown process (3) interaction of steam ejectors with the primary plume i.e. flow transients as well as probability of afterburning. In this abstract we discuss unsteady analyses of the engine shutdown process. However, the final paper will include analyses of a staged startup, drawdown of the engine test cell pressure, and risk assessment of potential afterburning in the facility. Unsteady simulations have been carried out to study the engine shutdown process in the facility and understand the physics behind the interactions between the steam ejectors, the test cell and the supersonic diffuser. As a first approximation, to understand the dominant unsteady mechanisms in the engine test cell and the supersonic diffuser, the turning duct in the facility was removed. As the engine loses power a rarefaction wave travels downstream that disrupts the shock cell structure in the supersonic diffuser. Flow from the test cell is seen to expand into the supersonic diffuser section and re-pressurizes the area around the nozzle along with a upstream traveling compression wave that emanates from near the first stage ejectors. Flow from the first stage ejector expands to the center of the duct and a new shock train is formed between the first and second stage ejectors. Both stage ejectors keep the facility pressurized and prevent any large amplitude pressure fluctuations from affecting the engine nozzle. The resultant pressure loads the nozzle experiences in the shutdown process are small.

Ahuja, Vineet↗

NASA NTRS2008

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively. The Mach Number distribution shows small yet distinct differences between the two cases such as locations of shocks/shock reflections and a slightly different impingement point on the wall of the diffuser from the expansion at the exit of the nozzle. Similarly the temperature distribution indicates different flow recirculation patterns in the test cell. Both cases capture all the essential flow phenomena such as the shock-boundary layer interaction, plume expansion, expansion of the first stage ejectors, mixing between the engine plume and the first stage ejector flow and pressurization due to the first stage ejectors. The final paper will discuss thermal loads on the walls of the diffuser and cooling mechanisms investigated.

Allgood, Daniel C.↗

NASA NTRS2009

NEXT Long-Duration Test After 11,570 h and 237 kg of Xenon Processed

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to validate and qualify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the mission-derived throughput requirement of 300 kg. This wear test is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of September 1, 2007, the thruster has accumulated 11,570 h of operation primarily at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. The thruster has processed 237 kg of xenon surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 (DS1) flight spare. The NEXT LDT has demonstrated a total impulse of 9.78 10(exp 6) N(dot)s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. Lifetime-limiting component erosion rates have been consistent with the NEXT service life assessment, which predicts the earliest failure sometime after 750 kg of xenon propellant throughput; well beyond the mission-derived lifetime requirement. The NEXT wear test data confirm that the erosion of the discharge keeper orifice, enlarging of nominal-current-density accelerator grid aperture cusps at full-power, and the decrease in cold grid-gap observed during NSTAR wear testing have been mitigated in the NEXT design. NEXT grid-gap data indicate a hot grid-gap at full-power that is 60 percent of the nominal cold grid-gap. This paper presents the status of the NEXT LDT to date with emphasis on comparison to the NSTAR extended life test results.

Soulas, George C.↗

NASA NTRS2008

Lunar Surface Access Module Descent Engine Turbopump Technology: Detailed Design

The need for a high specific impulse LOX/LH2 pump-fed lunar lander engine has been established by NASA for the new lunar exploration architecture. Studies indicate that a 4-engine cluster in the thrust range of 9,000-lbf each is a candidate configuration for the main propulsion of the manned lunar lander vehicle. The lander descent engine will be required to perform minor mid-course corrections, a Lunar Orbit Insertion (LOI) burn, a de-orbit burn, and the powered descent onto the lunar surface. In order to achieve the wide range of thrust required, the engines must be capable of throttling approximately 10:1. Working under internal research and development funding, NASA Marshall Space Flight Center (MSFC) has been conducting the development of a 9,000-lbf LOX/LH2 lunar lander descent engine testbed. This paper highlights the detailed design and analysis efforts to develop the lander engine Fuel Turbopump (FTP) whose operating speeds range from 30,000-rpm to 100,000-rpm. The capability of the FTP to operate across this wide range of speeds imposes several structural and dynamic challenges, and the small size of the FTP creates scaling and manufacturing challenges that are also addressed in this paper.

Alarez, Erika↗

NASA NTRS1940

Experimental Determination of Exhaust Gas Thrust, Special Report

This investigation presents the results of tests made on a radial engine to determine the thrust that can be obtained from the exhaust gas when discharged from separate stacks and when discharged from the collector ring with various discharge nozzles. The engine was provided with a propeller to absorb the power and was mounted on a test stand equipped with scales for measuring the thrust and engine torque. The results indicate that at full open throttle at sea level, for the engine tested, a gain in thrust horsepower of 18 percent using separate stacks, and 9.5 percent using a collector ring and discharge nozzle, can be expected at an air speed of 550 miles per hour.

Pinkel, Benjamin↗

NASA NTRS1939

Experiments on the Recovery of Waste Heat in Cooling Ducts, Special Report

Tests have been conducted in the N.A.C.A. full-scale wind tunnel to investigate the partial recovery of the heat energy which is apparently wasted in the cooling of aircraft engines. The results indicate that if the radiator is located in an expanded duct, a part of the energy lost in cooling is recovered; however, the energy recovery is not of practical importance up to airplane speeds of 400 miles per hour. Throttling of the duct flow occurs with heated radiators and must be considered in designing the duct outlets from data obtained with cold radiators in the ducts.

Silverstein, Abe↗

NASA NTRS1931

Estimated Effect of Ring Cowl on the Climb and Ceiling of an Airplane, Special Report

Although the application of a ring cowl to an airplane with an air-cooled engine increases the maximum L/D and the high speed to an appreciable extent, the performance in climb and ceiling is not increased as much as one would expect without analyzing the conditions. When a ring cowl is installed on an airplane, the propeller is set at a higher pitch to allow the engine to turn its rated r.p.m. at the increased high speed. V/nD is increased and the propeller efficiency at high speed is increased slightly. The ratio of r.p.m. at climbing speed, V(sub c) , to the r.p.m. at maximum speed, V (sub m) is dependent upon the ratio of V(sub c) to V(sub m). The increase in V(sub c) for all airplane with ring cowl i s not as great as the increase in V(sub m), so that the ratio V(sub c)/V(sub m) is less than for the airplane without ring. Consequently the r.p.m. and full throttle thrust power available are less at V(sub c) for the airplane with ring cowl and in spite of the increase in L/D due to the installation of the ring, the excess thrust power available for climbing is not appreciably changed. The same method of reasoning accounts for the small increase in absolute ceiling in spite of a large increase in L/D maximum.

Louden, F. A.↗

NASA NTRS2007

Thermal Development Test of the NEXT PM1 ION Engine

NASA's Evolutionary Xenon Thruster (NEXT) is a next-generation high-power ion thruster under development by NASA as a part of the In-Space Propulsion Technology Program. NEXT is designed for use on robotic exploration missions of the solar system using solar electric power. Potential mission destinations that could benefit from a NEXT Solar Electric Propulsion (SEP) system include inner planets, small bodies, and outer planets and their moons. This range of robotic exploration missions generally calls for ion propulsion systems with deep throttling capability and system input power ranging from 0.6 to 25 kW, as referenced to solar array output at 1 Astronomical Unit (AU). Thermal development testing of the NEXT prototype model 1 (PM1) was conducted at JPL to assist in developing and validating a thruster thermal model and assessing the thermal design margins. NEXT PM1 performance prior to, during and subsequent to thermal testing are presented. Test results are compared to the predicted hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.

NASA's Evolutionary Xenon Thruster (NEXT)↗

NASA NTRS1947

Investigation of the I-40 Jet-Propulsion Engine in the Cleveland Altitude Wind Tunnel. V - Operational Characteristics: Operational Characteristics - 5

An investigation has been conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of the I-40 jet-propulsion engine over a range of pressure altitudes from 10,000 to 50,000 feet and ram-pressure ratios from 1.00 to 1.76. Engine operational data were obtained with the engine in the standard configuration and with various modifications of the fuel system, the electrical system, and the combustion chambers. The effects of altitude and airspeed on operating speed range, starting, windmilli.ng, acceleration, speed regulation, cooling, and vibration of the standard and modified engines were determined, and damage to parts was noted. Maximum engine speed was obtainable at all altitudes and airspeeds wi th each fuel-control system investigated. The minimum idling speed was raised by increases in altitude and airspeed. The lowest minimum stable speeds were obtained with the standard configuration using 40-gallon nozzles with individual metering plugs. The engine was started normally at altitudes as high as 20,000 feet with all of the fuel systems and ignition combinations except one. Ignition at 70,000 feet was difficult and, although successful ignition occurred, acceleration was slow and usually characterized by excessive tail-pipe temperature. During windmilling investigations of the engine equipped with the standard fuel system, the engine could not be started at ram-pressure ratios of 1.1 to 1.7 at altitudes of 10,000, 20,000 and 30,000 feet. When equipped with the production barometric and Monarch 40-gallon nozzles, the engine accelerated in 12 seconds from an engine speed of 6000 rpm to 11,000 rpm at 20,000 feet and an average tail-pipe temperature of 11000 F. At the same altitude and temperature, all the engine configurations had approximately the same rate of acceleration. The Woodward governor produced the safest accelerations, inasmuch as it could be adjusted to automatically prevent acceleration blow out. The engine speed was held constant by the Woodward governor and the Edwards regulator during simulated dives and climbs at constant throttle position. The bearing cooling system was satisfactory at all altitudes and airspeeds. The engines operated without serious failure, although the exhaust cone, the tail pipe, and the airplane fuselage were damaged during altitude starts.

Golladay, Richard L.↗

NASA NTRS1946

Investigation of Three Design Modifications of the NACA Injection Impeller in an R-3350 Engine

An investigation was conducted to determine the effects of three design modifications of the original NACA injection impeller on the performance of an R-3350 engine. Different methods of injecting the fuel into the impeller air stream were studied and evaluated from the individual cylinder fuel-air ratios and the resulting cylinder temperatures. Each impeller was tested for a range of engine powers normally used in flight operation. The relatively simple design of the original injection impeller produced approximately the same mixture- and temperature-distribution characteristics as the modified impellers of more complex design. None of the modifications appreciably affected the manifold pressure, the combustion-air flow, nor the throttle angle required to maintain a given engine power,

Hickel, Robert O.↗

NASA NTRS1950

Preliminary Performance Data on Westinghouse Electronic Power Regulator Operating on J34-WE-32 Turbojet Engine in Altitude Wind Tunnel

The behavior of the Westinghouse electronic power regulator operating on a J34-WE-32 turbojet engine was investigated in the NACA Lewis altitude wind tunnel at the request of the Bureau of Aeronautics, Department of the Navy. The object of the program was to determine the, steady-state stability and transient characteristics of the engine under control at various altitudes and ram pressure ratios, without afterburning. Recordings of the response of the following parameters to step changes in power lever position throughout the available operating range of the engine were obtained; ram pressure ratio, compressor-discharge pressure, exhaust-nozzle area, engine speed, turbine-outlet temperature, fuel-valve position, jet thrust, air flow, turbine-discharge pressure, fuel flow, throttle position, and boost-pump pressure. Representative preliminary data showing the actual time response of these variables are presented. These data are presented in the form of reproductions of oscillographic traces.

Ketchum, James R.↗

NASA NTRS2007

Sensitivity Testing of the NSTAR Ion Thruster

During the Extended Life Test of the DS1 flight spare ion thruster, the engine was subjected to sensitvity testing in order to characterize the macroscopic dependence of discharge chamber sensitivity to a +\-3% vatiation in main flow, cathode flow and beam current, and to +\5% variation in beam and accelerator voltage, was determined for the minimum- (THO), half- (TH8) and full power (TH15) throttle levels. For each power level investigared, 16 high/low operating conditions were chosen to vary the flows, beam current, and grid voltages in in a matrix that mapped out the entire parameter space. The matrix of data generated was used to determine the partial derivative or senitivity of the dependent parameters--discharge voltage, discharge current, discharge loss, double-to-single-ion current ratio, and neutralizer-keeper voltage--to the variation in the independent parameters--main flow, cathode flow, beam current, and beam voltage. The sensititivities of each dependent parameter with respect to each independent parameter were determined using a least-square fit routine. Variation in these sensitivities with thruster runtime was recorded over the duration of the ELT, to detemine if discharge performance changed with thruster wear. Several key findings have been ascertained from the sensitivity testing. Discharge operation is most sensitve to changes in cathode flow and to a lesser degree main flow. The data also confirms that for the NSTAR configuration plasma production is limited by primary electron input due to the fixed neutral population. Key sensitivities along with their change with thruster wear (operating time) will be presented. In addition double ion content measurements with an ExB probe will also be presented to illustrate beam ion production and content sensitivity to the discharge chamber operating parameteres.

electric propulsion↗

NASA NTRS2009

Platform Precision Autopilot Overview and Mission Performance

The Platform Precision Autopilot is an instrument landing system-interfaced autopilot system, developed to enable an aircraft to repeatedly fly nearly the same trajectory hours, days, or weeks later. The Platform Precision Autopilot uses a novel design to interface with a NASA Gulfstream III jet by imitating the output of an instrument landing system approach. This technique minimizes, as much as possible, modifications to the baseline Gulfstream III jet and retains the safety features of the aircraft autopilot. The Platform Precision Autopilot requirement is to fly within a 5-m (16.4-ft) radius tube for distances to 200 km (108 nmi) in the presence of light turbulence for at least 90 percent of the time. This capability allows precise repeat-pass interferometry for the Unmanned Aerial Vehicle Synthetic Aperture Radar program, whose primary objective is to develop a miniaturized, polarimetric, L-band synthetic aperture radar. Precise navigation is achieved using an accurate differential global positioning system developed by the Jet Propulsion Laboratory. Flight-testing has demonstrated the ability of the Platform Precision Autopilot to control the aircraft within the specified tolerance greater than 90 percent of the time in the presence of aircraft system noise and nonlinearities, constant pilot throttle adjustments, and light turbulence.

Strovers, Brian K.↗

NASA NTRS2009

Sea-Level Flight Demonstration and Altitude Characterization of a LO2/LCH4 Based Accent Propulsion Lander

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Collins, Jacob↗

NASA NTRS2009

The NASA In-Space Propulsion Technology Project, Products, and Mission Applicability

The In-Space Propulsion Technology (ISPT) Project, funded by NASA s Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This overview provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical thrusters, and systems analysis tools. Aerocapture investments improved: guidance, navigation, and control models of blunt-body rigid aeroshells; atmospheric models for Earth, Titan, Mars, and Venus; and models for aerothermal effects. Investments in electric propulsion technologies focused on completing NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6 to 7 kW throttle-able gridded ion system. The project is also concluding its High Voltage Hall Accelerator (HiVHAC) mid-term product specifically designed for a low-cost electric propulsion option. The primary chemical propulsion investment is on the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. The project is also delivering products to assist technology infusion and quantify mission applicability and benefits through mission analysis and tools. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.

Anderson, David J.↗

NASA NTRS2009

Status of the NEXT Long-Duration Test After 23,300 Hours of Operation

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated in June 2005, to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg per thruster from mission analyses. The LDT is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of July 2009, the thruster has accumulated 23,300 h of operation with extensive durations at the following input powers: 6.9, 4.7, 1.1, and 0.5 kW. The thruster has processed 427 kg of xenon surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster and approaching the NEXT development qualification throughput goal. The NEXT LDT has demonstrated a total impulse of 16.0 10(exp 6) N/s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. The NSTAR first-failure mode, accelerator aperture erosion leading to electron backstreaming, has been mitigated in the NEXT design. The severe NSTAR discharge cathode assembly erosion has been mitigated by a graphite keeper in the NEXT thruster. Tracking of the NEXT first failure mode, charge-exchange ion impingement on the accelerator grid causing hexagonal groove erosion, is consistent with model predictions and indicates thruster life greater than or equal to 750 kg throughput. This paper presents the status, performance data, and wear characteristics of the NEXT LDT to date.

Herman, Daniel A.↗

NASA NTRS2010

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility: Steady Predictions - Part 1

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively.

Allgood, Daniel C.↗