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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 1,765 records · Page 98
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

RL-10 Based Combined Cycle For A Small Reusable Single-Stage-To-Orbit Launcher

This paper discusses a new application of the combined propulsion known as the KLIN(TM) cycle, consisting of a thermally integrated deeply cooled turbojet (DCTJ) and liquid rocket engine (LRE). If based on the RL10 rocket engine family, the KLIN (TM) cycle makes a small single-stage-to-orbit (SSTO) reusable launcher feasible and economically very attractive. Considered in this paper are the concept and parameters of a small SSTO reusable launch vehicle (RLV) powered by the KLIN (TM) cycle (sSSTO(TM)) launcher. Also discussed are the benefits of the small launcher, the reusability, and the combined cycle application. This paper shows the significant reduction of the gross take off weight (GTOW) and dry weight of the KLIN(TM) cycle-powered launcher compared to an all-rocket launcher.

Balepin, Vladimir↗
NASA NTRS1999

A Multi-Stage Continuous-Duty Adiabatic Demagnetization Refrigerator

The design for a multi-stage adiabatic demagnetization refrigerator (ADR) that can provide continuous cooling at very low temperatures is presented. The ADR is being developed for use in x-ray, IR and sub-millimeter space astronomy missions which will employ large format detector arrays operating at 50 mK and lower and which may dissipate up to 10 microwatts. It is also being designed to reject heat slowly to a relatively warm heat sink (in the 6-10 K range), so that future missions may use mechanical cryocoolers instead of liquid helium for pre-cooling. The continuous nature of the device gives it a much higher cooling power per unit mass, allowing it to be much smaller and lighter than existing ADRs with comparable performance. Design details are discussed along with prototype test results.

Shirron, P. J.↗
NASA NTRS1999

Solar Thermal Upper Stage Cryogen System Engineering Checkout Test

The Solar Thermal Upper Stage technology (STUSTD) program is a solar thermal propulsion technology program cooperatively sponsored by a Boeing led team and by NASA MSFC. A key element of its technology program is development of a liquid hydrogen (LH2) storage and supply system which employs multi-layer insulation, liquid acquisition devices, active and passive thermodynamic vent systems, and variable 40W tank heaters to reliably provide near constant pressure H2 to a solar thermal engine in the low-gravity of space operation. The LH2 storage and supply system is designed to operate as a passive, pressure fed supply system at a constant pressure of about 45 psia. During operation of the solar thermal engine over a small portion of the orbit the LH2 storage and supply system propulsively vents through the enjoy at a controlled flowrate. During the long coast portion of the orbit, the LH2 tank is locked up (unvented). Thus, all of the vented H2 flow is used in the engine for thrust and none is wastefully vented overboard. The key to managing the tank pressure and therefore the H2 flow to the engine is to manage and balance the energy flow into the LH2 tank with the MLI and tank heaters with the energy flow out of the LH2 tank through the vented H2 flow. A moderate scale (71 cu ft) LH2 storage and supply system was installed and insulated at the NASA MSFC Test Area 300. The operation of the system is described in this paper. The test program for the LH2 system consisted of two parts: 1) a series of engineering tests to characterize the performance of the various components in the system: and 2) a 30-day simulation of a complete LEO and GEO transfer mission. This paper describes the results of the engineering tests, and correlates these results with analytical models used to design future advanced Solar Orbit Transfer Vehicles.

Olsen, A. D↗
NASA NTRS1999

Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system and in the case of SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA); the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately determine insulation masses for a vehicle such as the one described above, the aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thickness that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass.

Alexander, Reginald Andrew↗
NASA NTRS1999

Reusable First Stage Flight Demonstrator

This report discusses the development of a reusable first stage flight demonstrator. The attributes of this demonstrator model is that it (1) use traceable technologies and subsystems to create an operational vehicle; (2) is affordable; (3) be scaleable to full scale vehicle. The demonstrator is to act as a technology test bed for reusable vehicle technology. The programs features both technical and schedule/cost risk reduction, which are outlined.

Mobley, T.↗
NASA NTRS1999

A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

Stanley, Thomas Troy↗
NASA NTRS1999

The Effect of Air Preheat at Atmospheric Pressure on the Formation of NO(x) in the Quick-Mix Sections of an Axially Staged Combustor

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor concept has been proposed to minimize the formation of nitrogen oxides (NO(x)) in gas turbine systems. The success of this combustor strategy is dependent upon the efficiency of the mixing section bridging the fuel-rich and fuel-lean stages. Note that although these results were obtained from an experiment designed to study an RQL mixer, the link between mixing and NOx signatures is considerably broader than this application, in that the need to understand this link exists in most advanced combustors. The experiment reported herein was designed to study the effects of inlet air temperature on NO(x) formation in a mixing section. The results indicate that NO(x) emission is increased for all preheated cases compared to non-preheated cases. When comparing the various mixing modules, the affect of jet penetration is important, as this determines where NO(x) concentrations peak, and affects overall NO(x) production. Although jet air comprises 70 percent of the total airflow, the impact that jet air preheat has on overall NO(x) emissions is small compared to preheating both main and jet air flow.

Vardakas, M. A.↗
NASA NTRS1997

Investigation of Boundary Layer Development in Multi-Stage Low Pressure

A review of recent literature has been made in the areas of unsteady measurement in turbomachinery and LP turbine design. This has included a detailed review of methods for extracting quantitative shear stress data from surface mounted hot-film arrays. Research goals were identified on the basis of the literature survey and following interaction with the GE and NASA contracts. Progress has also been made in the detailed planning of the research program. An array of hot-film sensors has been mounted at mid-span on a second stage nozzle blade ready for installation in the current turbine buildup. A new technique of applying the hot-film array using double sided adhesive is being tested. The temperature in the vicinity of the hot-films will be monitored using a thermocouple attached to the blade pressure surface. This will increase the accuracy of the data reduction process compared with the methods applied by Halstead et at. (1995) and others. Data obtained from this first set of sensors, along with data from the hot-wire investigation will help to determine the best locations for surface hot-film instrumentation for subsequent builds of the turbine. Other LSRT instrumentation issues which have been addressed include calibration of the torque measurement system and procurement of custom design hot-wire sensors required to compensate for the slope flow path. Preliminary attempts at modeling the boundary layer transition process have been completed. A steady, coupled viscous-inviscid flow solver (the MISES code of Drela and Youngren) was used to calculate the steady-flow boundary layer transition location.

Source record↗
NASA NTRS1999

A Collaborative Analysis Tool for Thermal Protection Systems for Single Stage to Orbit Launch Vehicles

Presented is a design tool and process that connects several disciplines which are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system and in the case of SSTO vehicles with air breathing propulsion, which is currently being studied by the National Aeronautics and Space Administration (NASA); the thermal protection system (TPS) is linked directly to almost every major system. The propulsion system pushes the vehicle to velocities on the order of 15 times the speed of sound in the atmosphere before pulling up to go to orbit which results high temperatures on the external surfaces of the vehicle. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. To adequately determine insulation masses for a vehicle such as the one described above, the aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thicknesses that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass. The problem described is an example of the need for collaborative design and analysis. Analysis tools are being developed to facilitate these collaborative efforts. RECIPE is a cross-platform application capable of hosting a number of engineers and designers across the Internet for distributed and collaborative engineering environments. Such integrated system design environments allow for collaborative team design analysis for performing individual or reduced team studies. The analysis tools mentioned earlier are commonly run on different platforms and are usually run by different people. To facilitate the larger number of potential runs that may need to be made, RECIPE connects the computer codes that calculate the trajectory data, heat rate data, and TPS masses so that the output from each tool is easily transferred to the model input files that need it. This methodology is being applied to solve launch vehicle thermal design problems to shorten the design cycle, and enable the project team to evaluate design options. Results will be presented indicating the effectiveness of this as a collaborative design tool.

Alexander, Reginald A.↗
NASA NTRS2000

A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC engine model. HYFIM performs the aerodynamic analysis of forebodies and inlet characteristics of RBCC powered SSTO launch vehicles. HYFIM is applicable to the analysis of the ramjet/scramjet engine operations modes (Mach 3-12), and provides estimates of parameters such as air capture area, shock-on-lip Mach number, design Mach number, compression ratio, etc., based on a basic geometry routine for modeling axisymmetric cones, 2-D wedge geometries. HYFIM also estimates the variation of shock layer properties normal to the forebody surface. The thermal protection system (TPS) is directly linked to determination of the vehicle moldline and the shaping of the trajectory. Thermal protection systems to maintain the structural integrity of the vehicle must be able to mitigate the heat transfer to the structure and be lightweight. Herein lies the interdependency, in that as the vehicle's speed increases, the TPS requirements are increased. And as TPS masses increase the effect on the propulsion system and all other systems is compounded. The need to analyze vehicle forebody and engine inlet is critical to be able to design the RBCC vehicle. To adequately determine insulation masses for an RBCC vehicle, the hypersonic aerodynamic environment and aeroheating loads must be calculated and the TPS thicknesses must be calculated for the entire vehicle. To accomplish this an ascent or reentry trajectory is obtained using the computer code Program to Optimize Simulated Trajectories (POST). The trajectory is then used to calculate the convective heat rates on several locations on the vehicles using the Miniature Version of the JA70 Aerodynamic Heating Computer Program (MINIVER). Once the heat rates are defined for each body point on the vehicle, then insulation thicknesses that are required to maintain the vehicle within structural limits are calculated using Systems Improved Numerical Differencing Analyzer (SINDA) models. If the TPS masses are too heavy for the performance of the vehicle the process may be repeated altering the trajectory or some other input to reduce the TPS mass. E-PSURBCC is an "engine performance" model and requires the specification of inlet air static temperature and pressure as well as Mach number (which it pulls from the HYFIM and POST trajectory files), and calculates the corresponding stagnation properties. The engine air flow path geometry includes inlet, a constant area section where the rocket is positioned, a subsonic diffuser, a constant area afterburner, and either a converging nozzle or a converging-diverging nozzle. The current capabilities of E-PSURBCC ejector and ramjet mode treatment indicated that various complex flow phenomena including multiple choking and internal shocks can occur for combinations of geometry/flow conditions. For a given input deck defining geometry/flow conditions, the program first goes through a series of checks to establish whether the input parameters are sound in terms of a solution path. If the vehicle/engine performance fails mission goals, the engineer is able to collaboratively alter the vehicle moldline to change aerodynamics, or trajectory, or some other input to achieve orbit. The problem described is an example of the need for collaborative design and analysis. RECIPE is a cross-platform application capable of hosting a number of engineers and designers across the Internet for distributed and collaborative engineering environments. Such integrated system design environments allow for collaborative team design analysis for performing individual or reduced team studies. To facilitate the larger number of potential runs that may need to be made, RECIPE connects the computer codes that calculate the trajectory data, aerodynamic data based on vehicle geometry, heat rate data, TPS masses, and vehicle and engine performance, so that the output from each tool is easily transferred to the model input files that need it.

Stanley, Thomas Troy↗
NASA NTRS2000

Flight Results of the Chandra X-ray Observatory Inertial Upper Stage Space Mission

Under contract to NASA, a specially configured version of the Boeing developed Inertial Upper Stage (IUS) booster was provided by Boeing to deliver NASA's 1.5 billion dollar Chandra X-Ray Observatory satellite into a highly elliptical transfer orbit from a Shuttle provided circular park orbit. Subsequently, the final orbit of the Chandra satellite was to be achieved using the Chandra Integral Propulsion System (IPS) through a series of IPS burns. On 23 July 1999 the Shuttle Columbia (STS-93) was launched with the IUS/Chandra stack in the Shuttle payload bay. Unfortunately, the Shuttle Orbiter was unexpectantly inserted into an off-nominal park orbit due to a Shuttle propulsion anomaly occurring during ascent. Following the IUS/Chandra on-orbit deployment from the Shuttle, at seven hours from liftoff, the flight proven IUS GN&C system successfully injected Chandra into the targeted transfer orbit, in spite of the off-nominal park orbit. This paper describes the IUS GN&C system, discusses the specific IUS GN&C mission data load development, analyses and testing for the Chandra mission, and concludes with a summary of flight results for the IUS part of the Chandra mission.

Tillotson, R.↗
NASA NTRS2000

A Multidisciplinary Performance Analysis of a Lifting-Body Single-Stage-to-Orbit Vehicle

Lockheed Martin Skunk Works (LMSW) is currently developing a single-stage-to-orbit reusable launch vehicle called VentureStar(TM) A team at NASA Langley Research Center participated with LMSW in the screening and evaluation of a number of early VentureStar(TM) configurations. The performance analyses that supported these initial studies were conducted to assess the effect of a lifting body shape, linear aerospike engine and metallic thermal protection system (TPS) on the weight and performance of the vehicle. These performance studies were performed in a multidisciplinary fashion that indirectly linked the trajectory optimization with weight estimation and aerothermal analysis tools. This approach was necessary to develop optimized ascent and entry trajectories that met all vehicle design constraints. Significant improvements in ascent performance were achieved when the vehicle flew a lifting trajectory and varied the engine mixture ratio during flight. Also, a considerable reduction in empty weight was possible by adjusting the total oxidizer-to-fuel and liftoff thrust-to-weight ratios. However, the optimal ascent flight profile had to be altered to ensure that the vehicle could be trimmed in pitch using only the flow diverting capability of the aerospike engine. Likewise, the optimal entry trajectory had to be tailored to meet TPS heating rate and transition constraints while satisfying a crossrange requirement.

Tartabini, Paul V.↗
NASA NTRS1992

Delta II Geotail -- 1st Stage and Solid Motor Booster Erection

The Geotail mission's goal was to investigate the structure and dynamics of the geomagnetic tail that extends on the nightside of the Earth. The launch date was July 24, 1992. This video shows the Delta II on the pad, being prepared for the launch. The first stage and the solid motor booster are shown being moved into place on the rocket.

Source record↗
NASA NTRS1993

A Monolithic W-Band Three-Stage LNA Using 0.1 Micron InAlAs/InGaAs/InP HEMT Technology

A monolithic W-band three-stage Low Noise Amplifiers (LNA) based on 0.1 um pseudomorphic InAlAs/InGaAs/InP (High Electron Mobility Transistors (HEMTs) has been developed. this LNA demonstrated a noise figure of 4.3 dB and an associated small signal gain of 19 dB at 100 GHz with a low dc power consumption of 20 mW. This is the best reported monolithic W-band LNA performance using InP-based HEMT technology and demonstrates the potential of InP HEMT technology for higher millimeter-wave applications.

Wang, H.↗
NASA NTRS2000

Hyper-X Stage Separation Wind Tunnel Test Program

NASA's Hyper-X research program was developed primarily to flight demonstrate a supersonic combustion ramjet engine, fully integrated with a forebody designed to tailor inlet flow conditions and a free expansion nozzle/afterbody to produce positive thrust at design flight conditions. With a point-designed propulsion system, the vehicle must depend upon some other means for boost to its design flight condition. Clean separation from this initial propulsion system stage within less than a second is critical to the success of the flight. This paper discusses the early planning activity, background, and chronology that developed the series of wind tunnel tests to support multi degree of freedom simulation of the separation process. Representative results from each series of tests are presented and issues and concerns during the process and current status will be highlighted.

Woods, W. C.↗