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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 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

NASA Propulsion Concept Studies and Risk Reduction Activities for Resource Prospector Lander

The trade study has led to the selection of propulsion concept with the lowest cost and net lowest risk -Government-owned, flight qualified components -Meet mission requirements although the configuration is not optimized. Risk reduction activities have provided an opportunity -Implement design improvements while development with the early-test approach. -Gain knowledge on the operation and identify operation limit -Data to anchor analytical models for future flight designs; The propulsion system cold flow tests series have provided valuable data for future design. -The pressure surge from the system priming and waterhammer within component operation limits. -Enable to optimize the ullage volume to reduce the propellant tank mass; RS-34 hot fire tests have successfully demonstrated of using the engines for the RP mission -No degradation of performance due to extended storage life of the hardware. -Enable to operate the engine for RP flight mission scenarios, outside of the qualification regime. -Provide extended data for the thermal and GNC designs. Significant progress has been made on NASA propulsion concept design and risk reductions for Resource Prospector lander.

Trinh, Huu P.↗
NASA NTRS1980

Curved flow wind tunnnel test of F-18 aircraft

The curved flow capability of a stability wind tunnel was used to investigate the lateral directional characteristics of an F-18 aircraft. The model is described and the procedures used to obtain and correct the data and a graphical presentation of the results are presented. The results include graphs of lateral directional derivatives versus sideslip or static plots, the lateral directional static stability derivatives versus angle of attack, and finally the lateral directional derivatives versus nondimensional yaw rate for different angles of attack and sideslip. Results are presented for several configurations including complete, complete without vertical tails, complete without horizontal tails, fuselage wing and fuselage alone. Each of these were tested with and without wing leading edge extensions.

Lutze, F. H.↗
NASA NTRS1985

Computation of three-dimensional shock-wave/turbulent boundary-layer interaction flows

Solutions of the Reynolds-averaged Navier-Stokes equations, employing a two-equation turbulence model, are presented and compared with measurements from a series of supersonic shock-wave/boundary-layer interaction experiments. The test flows include swept-compression corners and axisymmetric bodies with either skewed or segmented conical flares. The solutions correctly predict all major qualitative features of the flow field. However, for flow fields with large separated zones, significant quantitative differences are observed between the computed and experimental results. Possible reasons for these differences are discussed.

Horstman, C. C.↗
NASA NTRS2015

Joint Detect and Avoid Flight Testing

This presentation gives insight into a joint flight testing effort that included participation from NASA, Honeywell, and General Atomics. The presentation includes roles and responsibilities, test flow, and encounter requirements and summary.

flight tests↗
NASA NTRS1978

Design and preliminary results of a semitranspiration cooled (Lamilloy) liner for a high-pressure high-temperature combustor

A Lamilloy combustor liner was designed, fabricated and tested in a combustor at pressures up to 8 atmospheres. The liner was fabricated of a three layer Lamilloy structure and designed to replace a conventional step louver liner. The liner is to be used in a combustor that provides hot gases to a turbine cooling test facility at pressures up to 40 atmospheres. The Lamilloy liner was tested extensively at lower pressures and demonstrated lower metal temperatures than the conventional liner, while at the same time requiring about 40 percent less cooling air flow. Tests conducted at combustor exit temperatures in excess of 2200 K have not indicated any cooling or durability problems with the Lamilloy linear.

Wear, J. D.↗
NASA NTRS1978

Design and preliminary results of a semitranspiration cooled /Lamilloy/ liner for a high-pressure high-temperature combustor

A Lamilloy combustor liner has been designed, fabricated and tested in a combustor at pressures up to 8 atmospheres. The liner was fabricated of a three layer Lamilloy structure and designed to replace a conventional step-louver liner. The liner will be used in a combustor that provides hot gases to a turbine cooling test facility at pressures up to 40 atmospheres. The Lamilloy liner was tested extensively at lower pressures and demonstrated lower metal temperatures than the conventional liner, while at the same time requiring about 40 percent less cooling air flow. Tests conducted at combustor exit temperatures in excess of 2200 K have not indicated any cooling or durability problems with the Lamilloy liner.

Wear, J. D.↗
NASA NTRS1984

Experimental evaluation of an advanced Space Shuttle Main Engine hot-gas manifold design concept

The Space Shuttle Main Engine's hot gas manifold (HGM) has been the subject of an experimental study aimed at the establishment of an aerodynamic data base to support the development of an advanced, three-dimensional, fluid dynamic analysis computer model. The advanced HGM design used in the study demonstrated improved flow uniformity in the fuel-side turbine exit and transfer duct exit regions. Major modifications were incorporated in the HGM flow test article model, using two large transfer ducts on the fuel turbine side in place of the three small transfer ducts of the present design. The HGM flow field data were found to be essentially independent of Reynolds number over the range examined.

Pelaccio, D. G.↗
NASA NTRS1986

An evaluation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

The aerodynamic performance of leading-edge flaps on three delta double-delta wing planforms having aspect ratios of 1.75, 2.11, and 2.50, have been investigated experimentally. The wings were mounted on a generic fuselage without an inlet canopy, or a vertical tail. The Mach numbers of the flow over the wings were 1.60, 1.90 and 2.16. A primary set of full-span leading-edge flaps with similar root and tip chords were tested on each wing, and several alternate flap planforms were tested on the aspect ratio 1.75 wings. It is found that all leading edge geometries were effective in reducing drag lifting over the range of wing aspect ratios and Mach numbers tested. Greater flap performance was obtained when primary flaps were applied to the delta planform. In general, the primary flap geometry yielded better performance than the alternative geometries tested. Flow visualization techniques were found to be useful for identifying the beneficial effects of leading-edge flap deflection on flow separation as well as fuselage interference effects. Black and white photographs of the delta and double-delta planforms are provided.

Covell, P. F.↗
NASA NTRS1990

An annular gas seal analysis using empirical entrance and exit region friction factors

Wall shear stress results from stationary-rotor flow tests of five annular gas seals are used to develop entrance and exit region friction factor models. The friction factor models are used in a bulk-flow seal analysis which predicts leakage and rotor-dynamic coefficients. The predictions of the analysis are compared to experimental results and to the predictions of Nelson's analysis (1985). The comparisons are for smooth-rotor seals with smooth and honeycomb-stators. The present analysis predicts the destabilizing cross-coupled stiffness of a seal better than Nelson's analysis. Both analyses predict direct damping well and direct stiffness poorly.

Elrod, D. A.↗
NASA NTRS1999

Investigation of the Rocket Induced Flow Field in a Rectangular Duct

Rocket-Based Combined Cycle (RBCC) concepts attempt to improve the performance of launch vehicles at all points in the launch trajectory and make highly reusable launch vehicles a reality. The Aerojet Strutjet RBCC concept consists of a variable geometry duct with internal, vertical struts that functions in ducted rocket, ramjet, scramjet, and pure rocket modes. These struts have rocket and turbine exhaust nozzles imbedded within them. The rocket flows create an ejector effect with the ingested air at subsonic flight velocities. In ramjet and scramjet modes, the fuel rich nozzle flows react with the ingested air producing an afterburner effect. Under a NASA Marshall Space Flight Center contract, the UAH Propulsion Research Center (PRC) has designed and built a Strutjet simulation facility. A scale model of a single strut has been built and is undergoing cold-flow testing to investigate the mixing of the rocket and turbine exhausts with the ingested air. A complementary experimental program is also underway to examine the induced flow-field generated by rocket nozzles confined in a rectangular duct. Characterizing the induced flow behavior is critical to understanding and optimizing the performance of future Strutjet-based RBCC propulsion systems. The proposed paper will present results from the rocket induced flow investigation.

Landrum, D. Brian↗
NASA NTRS2016

Effects of Alternate Leading Edge Cutback on the Space Shuttle Main Engine Low Pressure Fuel Pump

A higher order cavitation oscillation observed in the SSME low pressure fuel pump has been eliminated in water flow testing of a modified subscale replica of the inducer. The low pressure pump was modified by removing the outboard sections of two opposing blades of the four-bladed inducer, blending the "cutback" regions into the blades at the leading edge and tip, and removing material on the suction sides to decrease the exposed leading edge thickness. The leading edge tips of the cutback blades were moved approximately 25 degrees from their previous locations, thereby increasing one blade to blade spacing, decreasing the second, while simultaneously moving the cutback tips downstream. The test was conducted in MSFC's inducer test loop at scaled operating conditions in degassed and filtered water. In addition to eliminating HOC across the entire scaled operating regime, rotating cavitation was suppressed while the range of both alternate blade and asymmetric cavitation were increased. These latter phenomena, and more significantly, the shifts between these cavitation modes also resulted in significant changes to the head coefficient at low cavitation numbers. Reverse flow was detected at a slightly larger flow coefficient with the cutback inducer and suction capability was reduced by approximately 1 velocity head at and above approximately 90% of the reference flow coefficient. These performance changes along with more intense reverse flow are consistent with poor flow area management and increased incidence in the cutback region. Although the test demonstrated that the inducer modification was successful at eliminating the higher order cavitation across the entire scaled operating regime, different, previously unobserved, cavitation oscillations were introduced and significant performance penalties were imposed.

Mulder, Andrew↗
NASA NTRS2004

High Temperature Propulsion System Structural Seals for Future Space Launch Vehicles

Durable, flexible sliding seals are required in advanced hypersonic engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures of 2000 to 2500 F. Current seal designs do not meet the demanding requirements for future engines, so NASA s Glenn Research Center is developing advanced seals and preloading devices to overcome these shortfalls. An advanced ceramic wafer seal design and two types of seal preloading devices were evaluated in a series of compression, scrub, and flow tests. Silicon nitride wafer seals survived 2000 in. 1000 cycles) of scrubbing at 1600 F against an Inconel 625 rub surface with no chips or signs of damage. Flow rates measured for the wafers before and after scrubbing were almost identical and were up to 32 times lower than those recorded for the best braided rope seal flow blockers. Canted coil springs and silicon nitride compression springs showed promise conceptually as potential seal preloading devices to help maintain seal resiliency. A finite element model of the canted coil spring revealed that it should be possible to produce a spring out of high temperature materials for applications at 2000+ F.

Dunlap, Patrick H., Jr.↗
NASA NTRS2003

High Temperature Propulsion System Structural Seals for Future Space Launch Vehicles

Durable, flexible sliding seals are required in advanced hypersonic engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures of 2000 to 2500 F. Current seal designs do not meet the demanding requirements for future engines, so NASA's Glenn Research Center is developing advanced seals and preloading devices to overcome these shortfalls. An advanced ceramic wafer seal design and two types of seal preloading devices were evaluated in a series of compression, scrub, and flow tests. Silicon nitride wafer seals survived 2000 in. (1000 cycles) of scrubbing at room temperature against an Inconel 625 rub surface with no chips or signs of damage. Flow rates measured for the wafers before and after scrubbing were almost identical and were much lower than those recorded for the best braided rope seal flow blockers. Canted coil springs and silicon nitride compression springs showed promise conceptually as potential seal preloading devices to help maintain seal resiliency. A finite element model of the canted coil spring revealed that it should be possible to produce a spring out of high temperature materials for applications at 2000+ F.

Dunlap, Patrick H., Jr.↗
NASA NTRS2004

Further Investigations of Hypersonic Engine Seals

Durable, flexible sliding seals are required in advanced hypersonic engines to seal the perimeters of movable engine ramps for efficient, safe operation in high heat flux environments at temperatures of 2000 to 2500 F. Current seal designs do not meet the demanding requirements for future engines, so NASA's Glenn Research Center is developing advanced seals and preloading devices to overcome these shortfalls. An advanced ceramic wafer seal design and two silicon nitride compression spring designs were evaluated in a series of compression, scrub, and flow tests. Silicon nitride wafer seals survived 2000 in. (50.8 m) of scrubbing at 2000 F against a silicon carbide rub surface with no chips or signs of damage. Flow rates measured for the wafers before and after scrubbing were almost identical and were up to 32 times lower than those recorded for the best braided rope seal flow blockers. Silicon nitride compression springs showed promise conceptually as potential seal preload devices to help maintain seal resiliency.

Dunlap, Patrick H., Jr.↗
NASA NTRS2002

Particle Size Effects on Flow Properties of PS304 Plasma Spray Feedstock Powder Blend

The effects of BaF2-CaF2 particle size and size distribution on PS304 feedstock powder flowability have been investigated. Angular BaF2-CaF2 eutectic powders were produced by comminution and classified by screening to obtain 38 to 45 microns 45 to 106 microns, 63 to 106 microns, 45 to 53 microns, 63 to 75 microns, and 90 to 106 microns particle size distributions. The fluorides were added incrementally from 0 to 10 wt% to the other powder constituents of the PS304 feedstock: nichrome, chromia, and silver powders. The flow rate of the powder blends decreased linearly with increasing concentration of the fluorides. Flow was degraded with decreasing BaF2-CaF2 particle size and with increasing BaF2-CaF2 particle size distribution. A semiempirical relationship is offered to describe the PS304 powder blend flow behavior. The Hausner Ratio confirmed the funnel flow test results, but was slightly less sensitive to differences in BaF2-CaF2 particle size and size distribution. These findings may have applicability to other powders that do not flow easily, such as ceramic powders.

Stanford, Malcolm K.↗
NASA NTRSDate not supplied

Transition Measurements on the SWiFT Model in the National Transonic Facility

Boundary layer transition locations on the Swept Wing Flow Test (SWiFT) Hybrid Wing Body model were measured in the National Transonic Facility (NTF) at the NASA Langley Research Center using several different methods. These methods include sublimating chemical flow visualization for a limited number of low Reynolds number conditions, as well as mean static pressure and dynamic pressure measurements. Results are presented for a range of conditions in the NTF, demonstrating Mach and Reynolds number effects. The transition front locations obtained from the different techniques agree well. A variation of the static pressure transition detection method is described in which continuous angle of attack polars are performed. This approach enables detection of the transition front at more challenging (high Reynolds number) conditions than is possible with sparser data from pitch-pause angle-of-attack polars. The trip dot effectiveness is also verified using all of the available diagnostics. Additionally, low Reynolds number results are compared with those acquired on the same model at the Aircraft Research Association (ARA) facility, and these results show good agreement overall, suggesting similar freestream flow quality conditions of the two facilities.

National Transonic Facility↗
NASA NTRSDate not supplied

Transition Measurements on the SWiFT Model in the National Transonic Facility

Boundary layer transition locations on the Swept Wing Flow Test (SWiFT) Hybrid Wing Body model were measured in the National Transonic Facility (NTF) at the NASA Langley Research Center using several different methods. These methods include sublimating chemical flow visualization for a limited number of low Reynolds number conditions, as well as mean static pressure and dynamic pressure measurements. Results are presented for a range of conditions in the NTF, demonstrating Mach and Reynolds number effects. The transition front locations obtained from the different techniques agree well. A variation of the static pressure transition detection method is described in which continuous angle of attack polars are performed. This approach enables detection of the transition front at more challenging (high Reynolds number) conditions than is possible with sparser data from pitch-pause angle-of-attack polars. The trip dot effectiveness is also verified using all of the available diagnostics. Additionally, low Reynolds number results are compared with those acquired on the same model at the Aircraft Research Association (ARA) facility, and these results show good agreement overall, suggesting similar freestream flow quality conditions of the two facilities.

National Transonic Facility↗
DOE OSTI2012

Uncertainty methodology for the strongly coupled physical phenomena associated with annular flow

Best-Estimate plus Uncertainty (BEPU) methods are slowly supplanting the use of deterministic analysis methods for thermal-hydraulic analyses. As the uncertainty methodologies evolve it is expected that, where both experimental techniques allow and data are available, there will be a shift to quantifying the uncertainty in increasingly more fundamental parameters. For example, for annular flow in a three-field analysis environment (vapor, liquid film, droplet), the driving parameters would be: a) film interfacial shear stress, b) droplet drag, c) droplet entrainment rate and d) droplet deposition rate. An improved annular flow modeling package was recently developed and implemented in an in-house version of the COBRA-TF best-estimate subchannel analysis tool (Lane, 2009). Significant improvement was observed in the code-to-data predictions of several steam-water annular flow tests following the implementation of this modeling package; however, to apply this model set in formal BEPU analysis requires uncertainty distributions to be determined. The unique aspect of annular flow, and the topic of the present work, is the strong coupling between the interfacial drag, entrainment and deposition phenomena. Ideally the uncertainty in each phenomenon would be isolated; however, the situation is further complicated by an inability to experimentally isolate and measure the individual rate processes (particularly entrainment rate), which results in available experimental data that are inherently integral in nature. This paper presents a methodology for isolating the individual physical phenomena of interest, to the extent that the currently available experimental data allow, and developing the corresponding uncertainty distributions for annular flow. (authors)

22 GENERAL STUDIES OF NUCLEAR REACTORS↗