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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 523 records · Page 29
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

No-vent fill pressurization tests using a cryogen simulant

The results are described of an experimental program which studied the performance of various no-vent fill techniques for tank-to-tank liquid transfer. The tests were performed using a cryogen simulant (Freon-114) and a test bed consisting of a multiple tank/plumbing network that enabled studies of a variety of different inlet flow and active mixing regimes. Several results and conclusions were drawn from the 26 transfer experiments comprising the program. Most notable was the significant improvement in fill performance (i.e., minimized fill time and maximized fill fraction) with increased agitation of the liquid surface. Another was the close correlation between measured condensation rates and those predicted by recent theories which express condensation as a function of turbulent eddy effects on the liquid surface. In most cases, test data exhibited strong agreement with an analytical model which accounts for tank heat transfer and thermodynamics in a 1 g environment.

Schmidt, G. R.↗
NASA NTRS1992

Calculation of installation effects within performance computer programs

Gasturbine engine components as compressors, burners and turbines are usually tested on rigs prior to installation into an engine. In the engine, the component behaviour is different due to a variety of reasons. The installation effects are caused by small geometrical differences due to nonrepresentative rig operating temperatures and pressures, by different gas properties and Reynolds numbers and by radial as well as circumferential temperature and pressure profiles at the inlet to the component. For highly accurate performance predictions these rig-to-engine effects have to be taken into account. Traditionally the term "installation" is also used for describing all the differences in engine operation and behaviour between toothed and aircraft. Intake and afterbody drag, power offtake and bleed as well as intake pressure losses and inlet flow distortion have significant impact on airflow, thrust, specific fuel consumption and compressor stability. Using modem performance synthesis programs all these effects can be simulated reasonably well.

J Kurzke↗
NASA NTRS1992

No-vent fill pressurization tests using a cryogen simulant

The results are described of an experimental program which studied the performance of various no-vent fill techniques for tank-to-tank liquid transfer. The tests were performed using a cryogen simulant (Freon-114) and a test bed consisting of a multiple tank/plumbing network that enabled studies of a variety of different inlet flow and active mixing regimes. Several results and conclusions were drawn from the 26 transfer experiments comprising the program. Most notable was the significant improvement in fill performance (i.e., minimized fill time and maximized fill fraction) with increased agitation of the liquid surface. Another was the close correlation between measured condensation rates and those predicted by recent theories which express condensation as a function of turbulent eddy effects on the liquid surface. In most cases, test data exhibited strong agreement with an analytical model which accounts for tank heat transfer and thermodynamics in a 1 g environment.

Schmidt, G.↗
NASA NTRS1992

Mixing characteristics of injector elements in liquid rocket engines - A computational study

A computational study has been performed to better understand the mixing characteristics of liquid rocket injector elements. Variations in injector geometry as well as differences in injector element inlet flow conditions are among the areas examined in the study. Most results involve the nonreactive mixing of gaseous fuel with gaseous oxidizer but preliminary results are included that involve the spray combustion of oxidizer droplets. The purpose of the study is to numerically predict flowfield behavior in individual injector elements to a high degree of accuracy and in doing so to determine how various injector element properties affect the flow.

Lohr, Jonathan C.↗
NASA NTRS1984

Ames 40 X 80/80 X 120 Foot Wind Tunnel: Turning Vanes Design

A number of different turning vanes are designed for the NASA Ames wind tunnel. Computer codes are used to design and analyze the turning vanes to insure that they comply with their individual constraints. The presentation is given in viewgraph format and displays pressure coefficients for the different turning vanes as well as loss coefficients versus inlet flow angles.

Sanz, J.↗
NASA NTRS1993

Engine/airframe integration for waverider cruise vehicles

Engine-airframe integrated hypersonic waveriders have been optimized previously for both cruising flight and accelerators using inverse design flowfields to provide desirable aerodynamic performance and inlet properties. With engine installation, cruisers retain the high L/D advantages of the basic waverider form. To provide an understanding of waverider optimizer trends, an analytical development is presented which includes the presentation of a simplified model of a generic waverider shape. The model is shown to have good agreement with computational results, and it is used to demonstrate engineering tradeoffs in L/D, volume, and volume efficiency. Based on the demonstrated performance trends, and considering the desirability of uniform inlet flow, an improved waverider shape for engine-airframe integration is proposed using a hybrid cone-wedge generating flowfield.

Lewis, Mark J.↗
NASA NTRS1994

Hypersonic vehicle model and control law development using H(infinity) and micron synthesis

The control system design for a Single Stage To Orbit (SSTO) air breathing vehicle will be central to a successful mission because a precise ascent trajectory will preserve narrow payload margins. The air breathing propulsion system requires the vehicle to fly roughly halfway around the Earth through atmospheric turbulence. The turbulence, the high sensitivity of the propulsion system to inlet flow conditions, the relatively large uncertainty of the parameters characterizing the vehicle, and continuous acceleration make the problem especially challenging. Adequate stability margins must be provided without sacrificing payload mass since payload margins are critical. Therefore, a multivariable control theory capable of explicitly including both uncertainty and performance is needed. The H(infinity) controller in general provides good robustness but can result in conservative solutions for practical problems involving structured uncertainty. Structured singular value mu framework for analysis and synthesis is potentially much less conservative and hence more appropriate for problems with tight margins. An SSTO control system requires: highly accurate tracking of velocity and altitude commands while limiting angle-of-attack oscillations, minimized control power usage, and a stabilized vehicle when atmospheric turbulence and system uncertainty are present. The controller designs using H(infinity) and mu-synthesis procedures were compared. An integrated flight/propulsion dynamic mathematical model of a conical accelerator vehicle was linearized as the vehicle accelerated through Mach 8. Vehicle acceleration through the selected flight condition gives rise to parametric variation that was modeled as a structured uncertainty. The mu-analysis approach was used in the frequency domain to conduct controller analysis and was confirmed by time history plots. Results demonstrate the inherent advantages of the mu framework for this class of problems.

Gregory, Irene M.↗
NASA NTRS1995

Effects of Initial Conditions on a Single Jet in Crossflow

An experimental investigation of the effects of jet inlet flow conditions has been conducted for the isothermal mixing of a single jet injected into a crossflow. Jet penetration and mixing was studied using planar Mie scattering to measure time-averaged jet mixture fraction distributions. The effects of 'passive' control methods such as jet 'tabs' and jet swirl are reported. Mixing effectiveness, determined using a spatial unmixedness parameter based on the variance of the mean jet concentration distributions, was compared to a baseline case of a round jet injected into a uniform crossflow. All results are compared at a jet-to-mainstream momentum-flux ratio of 8.5. In the near-field, the mixing rates are similar to, or less than, the baseline configuration using this measure of mixedness. None of the tested configurations appear to significantly augment mixing within a downstream distance of 3 diameters of an equivalent-area round orifice.

Liscinsky, D. S.↗
NASA NTRS1995

Space Station Water Processor Mostly Liquid Separator (MLS)

This report presents the results of the development testing conducted under this contract to the Space Station Water Processor (WP) Mostly Liquid Separator (MLS). The MLS units built and modified during this testing demonstrated acceptable air/water separation results in a variety of water conditions with inlet flow rates ranging from 60 - 960 LB/hr.

Lanzarone, Anthony↗
NASA NTRS1999

Assessment of Integrated Nozzle Performance

This presentation highlights the activities that researchers at the NASA Lewis Research Center (LeRC) have been and will be involved in to assess integrated nozzle performance. Three different test activities are discussed. First, the results of the Propulsion Airframe Integration for High Speed Research 1 (PAIHSR1) study are presented. The PAIHSR1 experiment was conducted in the LeRC 9 ft x l5 ft wind tunnel from December 1991 to January 1992. Second, an overview of the proposed Mixer/ejector Inlet Distortion Study (MIDIS-E) is presented. The objective of MIDIS-E is to assess the effects of applying discrete disturbances to the ejector inlet flow on the acoustic and aero-performance of a mixer/ejector nozzle. Finally, an overview of the High-Lift Engine Aero-acoustic Technology (HEAT) test is presented. The HEAT test is a cooperative effort between the propulsion system and high-lift device research communities to assess wing/nozzle integration effects. The experiment is scheduled for FY94 in the NASA Ames Research Center (ARC) 40 ft x 80 ft Low Speed Wind Tunnel (LSWT).

Lambert, H. 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.↗
NASA NTRS2000

Hyper-X Storage 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, William C.↗
NASA NTRS2002

Computational Investigation of the Performance and Back-Pressure Limits of a Hypersonic Inlet

A computational analysis of Mach 6.2 operation of a hypersonic inlet with rectangular-to-elliptical shape transition has been performed. The results of the computations are compared with experimental data for cases with and without a manually imposed back-pressure. While the no-back-pressure numerical solutions match the general trends of the data, certain features observed in the experiments did not appear in the computational solutions. The reasons for these discrepancies are discussed and possible remedies are suggested. Most importantly, however, the computational analysis increased the understanding of the consequences of certain aspects of the inlet design. This will enable the performance of future inlets of this class to be improved. Computational solutions with back-pressure under-estimated the back-pressure limit observed in the experiments, but did supply significant insight into the character of highly back-pressured inlet flows.

Smart, Michael K.↗
NASA NTRS2001

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 on 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 are highlighted.

Woods, William C.↗
NASA NTRS2006

Space Shuttle Main Engine Low Pressure Oxidizer Turbo-Pump Inducer Dynamic Environment Characterization through Water Model and Hot-Fire Testing

The Low Pressure Oxidizer Turbopump (LPOTP) inducer on the Block II configuration Space Shuttle Main Engine (SSME) experienced blade leading edge ripples during hot firing. This undesirable condition led to a minor redesign of the inducer blades. This resulted in the need to evaluate the performance and the dynamic environment of the redesign, relative to the current configuration, as part of the design acceptance process. Sub-scale water model tests of the two inducer configurations were performed, with emphasis on the dynamic environment due to cavitation induced vibrations. Water model tests were performed over a wide range of inlet flow coefficient and pressure conditions, representative of the scaled operating envelope of the Block II SSME, both in flight and in ground hot-fire tests, including all power levels. The water test hardware, facility set-up, type and placement of instrumentation, the scope of the test program, specific test objectives, data evaluation process and water test results that characterize and compare the two SSME LPOTP inducers are discussed. In addition, dynamic characteristics of the two water models were compared to hot fire data from specially instrumented ground tests. In general, good agreement between the water model and hot fire data was found, which confirms the value of water model testing for dynamic characterization of rocket engine turbomachinery.

Patrick Arellano↗
NASA NTRS2007

Effects of Unsteadiness Due to Wake Passing on Rotor Blade Heat Transfer

14. ABSTRACT In a gas turbine engine, the turbine rotor blades are buffeted by the wakes of the vanes located upstream. There is a transient effect from the passing of wakes on the blade heat transfer. This transient effect has been computed for a representative rotor by introducing a wake upstream via an unsteady inlet flow boundary condition, or "gust" condition. Two cases of turbulent flow and laminar flow with Reynolds numbers of 385,000 and 385 respectively were considered. For the turbulent flow case a quasi-steady calculation was also performed. The variation in the unsteady heat transfer coefficient was found to be as high as 120 percent of the mean. For the turbulent flow case a quasisteady calculation was also performed. The time mean of the unsteady heat transfer, the mean of the quasi-steady variations and the steady results agree reasonably well on all blade locations except for the turbulent results which differ near the leading edge. The quasi-steady heat transfer results do not agree with the instantaneous unsteady results, although the time-mean values are similar.

Ameri, Ali A.↗
NASA NTRS2009

Improvements in Modeling 90 degree Bleed Holes for Supersonic Inlets

The modeling of porous bleed regions as boundary conditions in computational fluid dynamics (CFD) simulations of supersonic inlet flows has been improved through a scaling of sonic flow coefficient data for 90deg bleed holes. The scaling removed the Mach number as a factor in computing the sonic flow coefficient and allowed the data to be fitted with a quadratic equation, with the only factor being the ratio of the plenum static pressure to the surface static pressure. The implementation of the bleed model into the Wind-US CFD flow solver was simplified by no longer requiring the evaluation of the flow properties at the boundary-layer edge. The quadratic equation can be extrapolated to allow the modeling of small amounts of blowing, which can exist when recirculation of the bleed flow occurs within the bleed region. The improved accuracy of the bleed model was demonstrated through CFD simulations of bleed regions on a flat plate in supersonic flow with and without an impinging oblique shock. The bleed model demonstrated good agreement with experimental data and three-dimensional CFD simulations of bleed holes.

Slater, John W.↗