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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 577 records · Page 32
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

High Voltage TAL Performance

The performance of a two-stage, anode layer Hall thruster was evaluated. Experiments were conducted in single and two-stage configurations. In single-stage configuration, the thruster was operated with discharge voltages ranging from 300 to 1700 V. Discharge specific impulses ranged from 1630 to 4140 sec. Thruster investigations were conducted with input power ranging from 1 to 8.7 kW, corresponding to power throttling of nearly 9: 1. An extensive two-stage performance map was generated. Data taken with total voltage (sum of discharge and accelerating voltage) constant revealed a decrease in thruster efficiency as the discharge voltage was increased. Anode specific impulse values were comparable in the single and two-stage configurations showing no strong advantage for two-stage operation.

Jacobson, David T.↗
NASA NTRS2001

Development and Lab-Scale Testing of a Gas Generator Hybrid Fuel in Support of the Hydrogen Peroxide Hybrid Upper Stage Program

As part of a NASA funded contract to develop and demonstrate a gas generator cycle hybrid rocket motor for upper stage space motor applications, the development and demonstration of a low sensitivity, high performance fuel composition was undertaken. The ultimate goal of the development program was to demonstrate successful hybrid operation (start, stop, throttling) of the fuel with high concentration (90+%) hydrogen peroxide. The formulation development and lab-scale testing of a simple DOT Class 1.4c gas generator propellant is described. Both forward injected center perforated and aft injected end burner hybrid combustion behavior were evaluated with gaseous oxygen and catalytically decomposed 90% hydrogen peroxide. Cross flow and static environments were found to yield profoundly different combustion behaviors, which were further governed by binder type, oxidizer level and, significantly, oxidizer particle size. Primary extinguishment was accomplished via manipulation of PDL behavior and oxidizer turndown, which is enhanced with the hydrogen peroxide system. Laboratory scale combustor results compared very well with 11-inch and 24-inch sub-scale test results with 90% hydrogen peroxide.

Lund, Gary K.↗
NASA NTRS1996

Comparison of Hover/Ground-Effect Characteristics for a Joint Strike Fighter Configuration

Hover and ground-effect tests were conducted with the Lockheed-Martin Large Scale Powered Model (LSPM) during June-November 1995 at the NASA Ames Research Center's Outdoor Aerodynamics Research Facility (OARF). This was done in support of the Joint Strike Fighter (JSF) Program being lead by the Department of Defense. The program was previously metered to as the Joint Advanced Strike Technology (JAST) Program. The tests at the OARF included: engine thrust calibrations out of ground effect, measurements of individual nozzle jet pressure decay characteristics, and jet-induced hover force and moment measurements in and out of ground effect. The engine calibrations provide data correlating propulsion system throttle and nozzle settings to thrust forces and moments for the bare fuselage with the wings, canards, and tails removed. This provides propulsive forces and moments while minimizing any of the effects due to the presence of the airframe. The engine calibrations were used later to determine thrust for hover testing at the OARF and for transition testing which took place in the NASA Ames 80 by 120 foot Wind Tunnel. The jet decay characteristics reflect the jet entrainment properties and are related to the aircraft suckdown characteristics. The JSF program provided the opportunity to obtain model scale effects using two models; one at small scale and one at large scale. Examples of data from these tests will be presented out of ground effect which will demonstrate the effect of scale. For one JSF configuration, two small-scale models and one large-scale model were tested and obtained different values for the lift loss out of ground effect. These differences were examined and are found to be largely dependent on the jet efflux characteristics. The jet-induced hover forces and moments are presented in and out of ground effect. For the hover testing the model is methodically built up from the bare fuselage, to include the wings, canards, and vertical tails, to determine the effects of each. Other components such as weapons bay doors, landing gear doors, and other lift improvement devices and their effect on the jet-induced forces and moments are examined.

Hange, Craig↗
NASA NTRS2002

Steady and Unsteady Simulations of the Flow in an Impeller/Diffuser Stage

SLI engine designs will require pumps to throttle over a wide flow range while maintaining high performance. Unsteadiness generated by impeller/diffuser interaction is one of the major factors affecting off-design performance. Initial unsteady simulations are completed for impeller/diffuser stage. The Corsair simulations will continue across a wide flow range and for inducer/impeller/diffuser combinations. Results of unsteady simulations are being used to guide and explore new designs.

Canabal, Francisco↗
NASA NTRS2001

Unshrouded Centrifugal Turbopump Impeller Design Methodology

Turbopump weight continues to be a dominant parameter in the trade space for reduction of engine weight. Space Shuttle Main Engine weight distribution indicates that the turbomachinery make up approximately 30% of the total engine weight. Weight reduction can be achieved through the reduction of envelope of the turbopump. Reduction in envelope relates to an increase in turbopump speed and an increase in impeller head coefficient. Speed can be increased until suction performance limits are achieved on the pump or due to alternate constraints the turbine or bearings limit speed. Once the speed of the turbopump is set the impeller tip speed sets the minimum head coefficient of the machine. To reduce impeller diameter the head coefficient must be increased. A significant limitation with increasing head coefficient is that the slope of the head-flow characteristic is affected and this can limit engine throttling range. Unshrouded impellers offer a design option for increased turbopump speed without increasing the impeller head coefficient. However, there are several issues with regard to using an unshrouded impeller: there is a pump performance penalty due to the front open face recirculation flow, there is a potential pump axial thrust problem from the unbalanced front open face and the back shroud face, and since test data is very limited for this configuration, there is uncertainty in the magnitude and phase of the rotordynamic forces due to the front impeller passage. The purpose of the paper is to discuss the design of an unshrouded impeller and to examine the hydrodynamic performance, axial thrust, and rotordynamic performance. The design methodology will also be discussed. This work will help provide some guidelines for unshrouded impeller design.

Prueger, George H.↗
NASA NTRS2002

Translation Optics for 30 cm Ion Engine Thrust Vector Control

Data were obtained from a 30 cm xenon ion thruster in which the accelerator grid was translated in the radial plane. The thruster was operated at three different throttle power levels, and the accelerator grid was incrementally translated in the X, Y, and azimuthal directions. Plume data was obtained downstream from the thruster using a Faraday probe mounted to a positioning system. Successive probe sweeps revealed variations in the plume direction. Thruster perveance, electron backstreaming limit, accelerator current, and plume deflection angle were taken at each power level, and for each accelerator grid position. Results showed that the thruster plume could easily be deflected up to six degrees without a prohibitive increase in accelerator impingement current. Results were similar in both X and Y direction.

Haag, Thomas↗
NASA NTRS2002

Experimental Evaluation of a Subscale Gaseous Hydrogen/gaseous Oxygen Coaxial Rocket Injector

The next generation reusable launch vehicle may utilize a Full-Flow Stage Combustion (FFSC) rocket engine cycle. One of the key technologies required is the development of an injector that uses gaseous oxygen and gaseous hydrogen as propellants. Gas-gas propellant injection provides an engine with increased stability margin over a range of throttle set points. This paper summarizes an injector design and testing effort that evaluated a coaxial rocket injector for use with gaseous oxygen and gaseous hydrogen propellants. A total of 19 hot-fire tests were conducted up to a chamber pressure of 1030 psia, over a range of 3.3 to 6.7 for injector element mixture ratio. Post-test condition of the hardware was also used to assess injector face cooling. Results show that high combustion performance levels could be achieved with gas-gas propellants and there were no problems with excessive face heating for the conditions tested.

Smith, Timothy D.↗
NASA NTRS2002

Optimal Aircraft Control Upset Recovery With and Without Component Failures

This paper treats the problem of recovering sustainable nondescending (safe) flight in a transport aircraft after one or more of its control effectors fail. Such recovery can be a challenging goal for many transport aircraft currently in the operational fleet for two reasons. First, they have very little redundancy in their means of generating control forces and moments. These aircraft have, as primary control surfaces, a single rudder and pairwise elevators and aileron/spoiler units that provide yaw, pitch, and roll moments with sufficient bandwidth to be used in stabilizing and maneuvering the airframe. Beyond this, throttling the engines can provide additional moments, but on a much slower time scale. Other aerodynamic surfaces, such as leading and trailing edge flaps, are only intended to be placed in a position and left, and are, hence, very slow-moving. Because of this, loss of a primary control surface strongly degrades the controllability of the vehicle, particularly when the failed effector becomes stuck in a non-neutral position where it exerts a disturbance moment that must be countered by the remaining operating effectors. The second challenge in recovering safe flight is that these vehicles are not agile, nor can they tolerate large accelerations. This is of special importance when, at the outset of the recovery maneuver, the aircraft is flying toward the ground, as is frequently the case when there are major control hardware failures. Recovery of safe flight is examined in this paper in the context of trajectory optimization. For a particular transport aircraft, and a failure scenario inspired by an historical air disaster, recovery scenarios are calculated with and without control surface failures, to bring the aircraft to safe flight from the adverse flight condition that it had assumed, apparently as a result of contact with a vortex from a larger aircraft's wake. An effort has been made to represent relevant airframe dynamics, acceleration limits, and actuator limits faithfully, since these contribute to the lack of agility and control power that plays an important role in defining what can be achieved with the vehicle when it is in extremis.

Sparks, Dean W.↗
NASA NTRS2003

Solar Energetic Particle Variations

In the largest solar energetic-particle (SEP) events, acceleration occurs at shock waves driven out from the Sun by coronal mass ejections (CMEs). In fact, the highest proton intensities directly measured near Earth at energies up to approximately 1 GeV occur at the time of passage of shocks, which arrive about a day after the CMEs leave the Sun. CME-driven shocks expanding across magnetic fields can fill over half of the heliosphere with SEPs. Proton-generated Alfven waves trap particles near the shock for efficient acceleration but also throttle the intensities at Earth to the streaming limit early in the events. At high energies, particles begin to leak from the shock and the spectrum rolls downward to form an energy-spectral 'knee' that can vary in energy from approximately 1 MeV to approximately 1 GeV in different events. All of these factors affect the radiation dose as a function of depth and latitude in the Earth's atmosphere and the risk to astronauts and equipment in space. SEP ionization of the polar atmosphere produces nitrates that precipitate to become trapped in the polar ice. Observations of nitrate deposits in ice cores reveal individual large SEP events and extend back approximately 400 years. Unlike sunspots, SEP events follow the approximately 80-100-year Gleissberg cycle rather faithfully and are now at a minimum in that cycle. The largest SEP event in the last 400 years appears to be related to the flare observed by Carrington in 1859, but the probability of SEP events with such large fluences falls off sharply because of the streaming limit.

Reames, D. V.↗
NASA NTRS2003

Factors Influencing Solar Electric Propulsion Vehicle Payload Delivery for Outer Planet Missions

Systems analyses were performed for missions utilizing solar electric propulsion systems to deliver payloads to outer-planet destinations. A range of mission and systems factors and their affect on the delivery capability of the solar electric propulsion system was examined. The effect of varying the destination, the trip time, the launch vehicle, and gravity-assist boundary conditions was investigated. In addition, the affects of selecting propulsion system and power systems characteristics (including primary array power variation, number of thrusters, thruster throttling mode, and thruster Isp) on delivered payload was examined.

Cupples, Michael↗
NASA NTRS2003

Analysis of Parallel Burn, No-Crossfeed TSTO RLV Architectures and Comparison to Parallel Burn with Crossfeed and Series Burn Architectures

Three dominant Two Stage To Orbit (TSTO) class architectures were studied: Series Burn (SB), Parallel Bum with crossfeed (PBw/cf), and Parallel Burn, no-crossfeed (PBncf). The study goal was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or a SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a fixed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled booster and orbiter (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study observations were: 1) A PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) A PBncf TSTO architecture is feasible for systems that stage at mach 7. 2a) HH architectures can achieve a mass growth relative to PBw/cf of <20%. 2b) KH architectures can achieve a mass growth relative to Series Burn of <20%. 3) Center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter specific thrust to weight ratio and to the position of the orbiter required to align the nozzle heights at liftoff. 4) Thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 5) Performance for HH vehicles was better when staged at mach 7 instead of mach 5. The study suggests possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements.

Smith, Garrett↗
NASA NTRS2003

Test Rig for Evaluating Active Turbine Blade Tip Clearance Control Concepts

Improved blade tip sealing in the high pressure compressor and high pressure turbine can provide dramatic improvements in specific fuel consumption, time-on-wing, compressor stall margin and engine efficiency as well as increased payload and mission range capabilities of both military and commercial gas turbine engines. The preliminary design of a mechanically actuated active clearance control (ACC) system for turbine blade tip clearance management is presented along with the design of a bench top test rig in which the system is to be evaluated. The ACC system utilizes mechanically actuated seal carrier segments and clearance measurement feedback to provide fast and precise active clearance control throughout engine operation. The purpose of this active clearance control system is to improve upon current case cooling methods. These systems have relatively slow response and do not use clearance measurement, thereby forcing cold build clearances to set the minimum clearances at extreme operating conditions (e.g., takeoff, re-burst) and not allowing cruise clearances to be minimized due to the possibility of throttle transients (e.g., step change in altitude). The active turbine blade tip clearance control system design presented herein will be evaluated to ensure that proper response and positional accuracy is achievable under simulated high-pressure turbine conditions. The test rig will simulate proper seal carrier pressure and temperature loading as well as the magnitudes and rates of blade tip clearance changes of an actual gas turbine engine. The results of these evaluations will be presented in future works.

Lattime, Scott B.↗
NASA NTRS2003

Implementation of a Water Flow Control System into the ISS'S Planned Fluids & Combustion Facility

The Fluids and Combustion Facility (FCF) will become an ISS facility capable of performing basic combustion and fluids research. The facility consists of two independent payload racks specifically configured to support multiple experiments over the life of the ISS. Both racks will depend upon the ISS's Moderate Temperature Loop (MTL) for removing waste heat generated by the avionics and experiments operating within the racks. By using the MTL, constraints are imposed by the ISS vehicle on how the coolant resource is used. On the other hand, the FCF depends upon effective thermal control for maximizing life of the hardware and for supplying proper boundary conditions for the experiments. In the implementation of a design solution, significant factors in the selection of the hardware included ability to measure and control relatively low flow rates, ability to throttle flow within the time constraints of the ISS MTL, conserve energy usage, observe low mass and small volume requirements. An additional factor in the final design solution selection was considering how the system would respond to a loss of power event. This paper describes the method selected to satisfy the FCF design requirements while maintaining the constraints applied by the ISS vehicle.

Edwards, Daryl A.↗
NASA NTRS2003

Status of the NEXT Ion Engine Wear Test

The status of the NEXT 2000 hour wear test is presented. This test is being conducted with a 40 cm engineering model ion engine, designated EM1, at a beam current higher than listed on the NEXT throttle table. Pretest performance assessments demonstrated that EM1 satisfies all thruster performance requirements. As of 7/3/03, the ion engine has accumulated 406 hours of operation at a thruster input power of 6.9 kW. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, has been steady to date with no indications of performance degradation. Images of the downstream discharge cathode, neutralizer, and accelerator aperture surfaces have exhibited no significant erosion to date.

Soulas, George C.↗
NASA NTRS2003

Adaptive Gas Turbine Engine Control for Deterioration Compensation Due to Aging

This paper presents an ad hoc adaptive, multivariable controller tuning rule that compensates for a thrust response variation in an engine whose performance has been degraded though use and wear. The upset appears when a large throttle transient is performed such that the engine controller switches from low-speed to high-speed mode. A relationship was observed between the level of engine degradation and the overshoot in engine temperature ratio, which was determined to cause the thrust response variation. This relationship was used to adapt the controller. The method is shown to work very well up to the operability limits of the engine. Additionally, since the level of degradation can be estimated from sensor data, it would be feasible to implement the adaptive control algorithm on-line.

Litt, Jonathan S.↗
NASA NTRS2002

Innovative, High-Pressure, Cryogenic Control Valve: Short Face-to-Face, Reduced Cost

A control valve that can throttle high-pressure cryogenic fluid embodies several design features that distinguish it over conventional valves designed for similar applications. Field and design engineers worked together to create a valve that would simplify installation, trim changes, and maintenance, thus reducing overall cost. The seals and plug stem packing were designed to perform optimally in cryogenic temperature ranges. Unlike conventional high-pressure cryogenic valves, the trim size can be changed independent of the body. The design feature that provides flexibility for changing the trim is a split body. The body is divided into an upper and a lower section with the seat ring sandwiched in between. In order to maintain the plug stem packing at an acceptable sealing temperature during cryogenic service, heat-exchanging fins were added to the upper body section. The body is made of stainless steel. The seat ring is made of a nickel-based alloy having a coefficient of thermal expansion less than that of the body material. Consequently, when the interior of the valve is cooled cryogenically, the body surrounding the seat ring contracts more than the seat ring. This feature prevents external leakage at the body-seat joint. The seat ring has been machined to have small, raised-face sealing surfaces on both sides of the seal groove. These sealing surfaces concentrate the body bolt load over a small area, thereby preventing external leakage. The design of the body bolt circle is different from that of conventional highpressure control valves. Half of the bolts clamp the split body together from the top, and half from the bottom side. This bolt-circle design allows a short, clean flow path, which minimizes frictional flow losses. This bolt-circle design also makes it possible to shorten the face-toface length of the valve, which is 25.5 in. (65 cm). In contrast, a conventional, high-pressure control valve face-to-face dimension may be greater than 40 in. (>1 m) long.

Source record↗
NASA NTRS2003

Space Shuttle Main Engine Implications for the Abort-to-Orbit Off-the-Pad Study

In 2001, the Space Shuttle Main Engine (SSME) project office was contacted by the Space Shuttle Ascent Guidance, Navigation, and Control group to provide the engine perspective for an Abort-to-Orbit (ATO) study. The purpose of the AT0 Off-the-Pad study was to determine the feasibility of eliminating the Return to Launch Site and Transatlantic abort modes by using a five-segment solid rocket booster and throttling the remaining SSMEs to a higher power level. This would enable all abort modes to be Abort-to-Orbit. The SSME project office at Marshall Space Flight Center collaborated with MSFC's Space Transportation Directorate and Rocketdyne Propulsion and Power to provide the AT0 Off-the-Pad study with the analysis required. Power levels at 109%, 1 11% and 1 13% of rated power level were studied as well as mixture ratio decreases down to 5.85. SSME was to evaluate and define the technical and programmatic impacts to certify the SSME to these abort power levels. The SSME systems analysis group performed a steady state analysis using the SSME power balance model to determine if there were any technical issues associated with higher power level, low mixture ratio operation. Based on each power level/mixture ratio combination, an engine certification plan was created and a preliminary probabilistic risk assessment was performed. The results showed favorable results for higher power 1evel/lower mixture ratio SSME operation. In nearly all performance and redline parameters, the traded engine operation was encompassed by nominal engine performance of a prior engine configuration.

Schoffstoll, Dayna L.↗
NASA NTRS2003

Experimental Investigation of Centrifugal Compressor Stabilization Techniques

Results from a series of experiments to investigate techniques for extending the stable flow range of a centrifugal compressor are reported. The research was conducted in a high-speed centrifugal compressor at the NASA Glenn Research Center. The stabilizing effect of steadily flowing air-streams injected into the vaneless region of a vane-island diffuser through the shroud surface is described. Parametric variations of injection angle, injection flow rate, number of injectors, injector spacing, and injection versus bleed were investigated for a range of impeller speeds and tip clearances. Both the compressor discharge and an external source were used for the injection air supply. The stabilizing effect of flow obstructions created by tubes that were inserted into the diffuser vaneless space through the shroud was also investigated. Tube immersion into the vaneless space was varied in the flow obstruction experiments. Results from testing done at impeller design speed and tip clearance are presented. Surge margin improved by 1.7 points using injection air that was supplied from within the compressor. Externally supplied injection air was used to return the compressor to stable operation after being throttled into surge. The tubes, which were capped to prevent mass flux, provided 9.3 points of additional surge margin over the baseline surge margin of 11.7 points.

Skoch, Gary J.↗