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At least 73 records · Page 4

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NASA NTRS1991

Shuttle-C auxiliary propulsion system certification approach and low-cost evolution concepts

Shuttle-C provides NASA with a heavy-lift cargo capability that is a low-cost derivative of the current Space Shuttle. This system can deliver 80,000 to 140,000 pounds of payload to low earth orbit within 4 years of authority to proceed. Shuttle-C will share th existing Shuttle launch facilities at Kennedy Space Center (KSC). The Shuttle-C vehicle incorporates an auxiliary propulsion system (APS) which provides separation from the external tank (ET), orbit circularization, on-orbit attitude hold, and deorbit capability. The APS is derived from the Shuttle orbiter aft reaction control system (ARCS) and orbital maneuvering system (OMS) with the objective of achieving minimum life cycle cost for this expendable vehicle. The resulting APS system will be described, its plan for a minimum-cost flight certification utilizing a cold-flow test program explained, and low-cost evolution by component design modification and part substitution presented.

Harsh, M. George↗

NASA NTRS1991

Experimental investigation of the flow field in the head-end star slot section of a solid rocket motor

A test program has been developed to investigate the complex flowfield in the head-end, star grain section of a solid rocket motor. Documentation of the flowfield pattern is to be obtained using both flow visualization techniques and quantitative methods, including static pressure measurements, heat transfer coefficient measurements, and hot-wire anemometry. A cold-flow, one-tenth-scale model based on the geometry of the Space Shuttle SRM head-end section and its single-port igniter, along with two models of a four-port igniter, have been constructed. The scale factor is derived from a scaling analysis which matches Reynolds number between the model and the full-scale flowfield in the star grain section. The NASA Marshall Space Flight Center 14 x 14 inch trisonic wind tunnel will be utilized for the cold flow tests. Since the work is still in progress, no results are included.

Ciucci, A.↗

NASA NTRS1993

A semi-empirical model for heat transfer and flame spreading in slotted rocket motors

This paper presents analytical and experimental results for the cold-flow aerothermodynamics of the head-end star slot region of the Space Shuttle solid rocket motor (SRM). Particular attention is paid to the determination of the heat transfer from the developing igniter flow to the propellant grain (slot wall). Igniter geometries considered include a single port configuration with flow along the axis of the motor (present Shuttle configuration) and a multiport 45-deg canted nozzle configuration which is similar (but not identical) to the proposed ASRM igniter geometry.

Jenkins, R. M.↗

NASA NTRS1957

Comparison of experimental with theoretical total-pressure loss in parallel-walled turbojet combustors

An experimental investigation of combustor total-pressure loss was undertaken to confirm previous theoretical analyses of effects of geometric and flow variables and of heat addition. The results indicate that a reasonable estimate of cold-flow total-pressure-loss coefficient may be obtained from the theoretical analyses. Calculated total-pressure loss due to heat addition agreed with experimental data only when there was no flame ejection from the liner at the upstream air-entry holes.

Dittrich, Ralph T↗

NASA NTRS1956

A Theory for Stability and Buzz Pulsation Amplitude in Ram Jets and an Experimental Investigation Including Scale Effects

From a theory developed on a quasi-one-dimensional-flow basis, it is found that the stability of the ram jet is dependent upon the instantaneous values of mass flow and total pressure recovery of the supersonic diffuser and immediate neighboring subsonic diffuser. Conditions for stable and unstable flow are presented. The theory developed in the report is in agreement with the experimental data of NACA-TN-3506 and NACA-RM-L50K30. A simple theory for predicting the approximate amplitude of small pressure pulsation in terms of mass-flow decrement from minimum-stable mass flow is developed and found to agree with experiments. Cold-flow tests at a Mach number of 1.94 of ram-jet models having scale factors of 3.15:1 and Reynolds number ratios of 4.75:1 with several supersonic diffuser configurations showed only small variations in performance between geometrically similar models. The predominant variation in steady-flow performance resulted from the larger boundary layer in the combustion chamber of the low Reynolds number models. The conditions at which buzz originated were nearly the same for the same supersonic diffuser (cowling-position angle) configurations in both large and small diameter models. There was no appreciable variation in stability limits of any of the models when the combustion-chamber length was increased by a factor of three. The unsteady-flow performance and wave patterns were also similar when considered on a reduced-frequency basis determined from the relative lengths of the model. The negligible effect of Reynolds number on stability of the off-design configurations was not anticipated in view of the importance of boundary layer to stability, and this result should not be construed to be generally applicable. (author)

Trimpi, Robert L↗

NASA NTRS1993

SSME fuel preburner injector characterization

A project has been initiated at the Marshall Space Flight Center to determine if preburner inter- or intra-element mixture ratio maldistributions are the cause of temperature variations in the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) turbine inlet region. Temperature nonuniformity may contribute to the many problems experienced in this region. The project will involve high pressure cold-flow testing and Computational Fluid Dynamics (CFD) modeling.

Hutt, John J.↗

NASA NTRS1994

Design and aerodynamic performance evaluation of a high-work mixed flow turbine stage

As axial and radial turbine designs have been pushed to their aerothermodynamic and mechanical limits, the mixed-flow turbine (MFT) concept has been projected to offer performance and durability improvements, especially when ceramic materials are considered. The objective of this NASA/U.S. Army sponsored mixed-flow turbine (AMFT) program was to determine the level of performance attainable with MFT technology within the mechanical constraints of 1997 projected ceramic material properties. The MFT geometry is similar to a radial turbine, exhibiting a large radius change from inlet to exit, but differing in that the inlet flowpath is not purely radial, nor axial, but mixed; it is the inlet geometry that gives rise to the name 'mixed-flow'. The 'mixed' orientation of the turbine inlet offers several advantages over radial designs by allowing a nonzero inlet blade angle yet maintaining radial-element blades. The oblique inlet not only improves the particle-impact survivability of the design, but improves the aerodynamic performance by reducing the incidence at the blade inlet. The difficulty, however, of using mixed-flow geometry lies in the scarcity of detailed data and documented design experience. This paper reports the design of a MFT stage designed with the intent to maximize aerodynamic performance by optimizing design parameters such as stage reaction, rotor incidence, flowpath shape, blade shape, vane geometry, and airfoil counts using 2-D, 3-D inviscid, and 3-D viscous computational fluid dynamics code. The aerodynamic optimization was accomplished while maintaining mechanical integrity with respect to vibration and stress levels in the rotor. A full-scale cold-flow rig test was performed with metallic hardware fabricated to the specifications of the hot ceramic geometry to evaluate the stage performance.

Neri, Remo N.↗

NASA NTRS1994

Analytical and experimental studies of impinging liquid jets

Impinging injectors are a common type of injector used in liquid propellant rocket engines and are typically used in engines where both propellants are injected as a liquid, e.g., engines using LOX/hydrocarbon and storable propellant combinations. The present research program is focused on providing the requisite fundamental understanding associated with impinging jet injectors for the development of an advanced a priori combustion stability design analysis capability. To date, a systematic study of the atomization characteristics of impinging liquid jets under cold-flow conditions have been completed. Effects of orifice diameter, impingement angle, pre-impingement length, orifice length-to-diameter ratio, fabrication procedure, jet flow condition and jet velocity under steady and oscillating, and atmospheric- and high-pressure environments have been investigated. Results of these experimental studies have been compared to current models of sheet breakup and drop formation. In addition, the research findings have been scrutinized to provide a fundamental explanation for a proven empirical correlation used in the design of stable impinging injector-based rocket engines.

Ryan, H. M.↗

NASA NTRS1994

Twin engine afterbody model

This test was originally conducted to determine the effects of several empennage and afterbody parameters on the aft-end aerodynamic characteristics of a twin-engine fighter-type configuration. Model variables were as follows: horizontal tail axial location and incidence, vertical tail axial location and configuration (twin-vs single-tail arrangements), tail booms, and nozzle power setting. Jet propulsion was simulated by exhausting high-pressure, cold-flow air from the nozzles. Following a successful test conducted on a single engine nacelle model to validate a CFD code, this model was chosen to be instrumented with pressure taps on the afterbody and nozzles and used as a follow-on test, providing a more complex geometry for the CFD code validation. A more limited test matrix was run to collect the pressure data, employing only the twin-tail configuration and varying only the horizontal and vertical tail locations. Mach number was varied from 0.6 to 1.2. Nozzle pressure ratio was varied from jet-off to 8. Angle-of-attack varied from 0 to 8 deg.

Wing, David J.↗

NASA NTRS1996

Turbulent Premixed Flames in Microgravity

The experimental cold-flow facility is now full operational and is currently being used to obtain baseline turbulence data in a Couette flow. The baseline turbulence data is necessary to confirm the capability of the chosen device to generate and maintain the required turbulence intensity. Subsequent reacting flow studies will assume that a similar turbulent flow field exists ahead of the premixed flame. Some modifications and refinements had to be made to enable accurate measurements. It consists of two rollers, one (driven by a motor) which drives a continuous belt and four smaller rollers used to set the belt spacing and tension to minimize belt flutter. The entire assemble is enclosed in a structure that has the dimensions to enable future drop tower experiments of the hot facility. All critical dimensions are the same as the original plans except for the pulley ratio which has been changed to enable a wider operating regime in terms of the Reynolds number. With the current setup, Reynolds numbers as low as 100 and as high as 14,000 can be achieved. This is because the in-between belt spacing can be varied from 1 cm to 7.6 cm, and the belt speed can be accurately varied from .15 m/sec to 3.1 m/sec.

Menon, Suresh↗

NASA NTRS1997

An Experimental Study of Swirling Flows as Applied to Annular Combustors

This thesis presents an experimental study of swirling flows with direct applications to gas turbine combustors. Two separate flowfields were investigated: a round, swirling jet and a non-combusting annular combustor model. These studies were intended to allow both a further understanding of the behavior of general swirling flow characteristics, such as the recirculation zone, as well as to provide a base for the development of computational models. In order to determine the characteristics of swirling flows the concentration fields of a round, swirling jet were analyzed for varying amount of swirl. The experimental method used was a light scattering concentration measurement technique known as marker nephelometry. Results indicated the formation of a zone of recirculating fluid for swirl ratios (rotational speed x jet radius over mass average axial velocity) above a certain critical value. The size of this recirculation zone, as well as the spread angle of the jet, was found to increase with increase in the amount of applied swirl. The annular combustor model flowfield simulated the cold-flow characteristics of typical current annular combustors: swirl, recirculation, primary air cross jets and high levels of turbulence. The measurements in the combustor model made by the Laser Doppler Velocimetry technique, allowed the evaluation of the mean and rms velocities in the three coordinate directions, one Reynold's shear stress component and the turbulence kinetic energy: The primary cross jets were found to have a very strong effect on both the mean and turbulence flowfields. These cross jets, along with a large step change in area and wall jet inlet flow pattern, reduced the overall swirl in the test section to negligible levels. The formation of the strong recirculation zone is due mainly to the cross jets and the large step change in area. The cross jets were also found to drive a four-celled vortex-type motion (parallel to the combustor longitudinal axis) near the cross jet injection plane.

Seal, Michael Damian, II↗

NASA NTRS1996

Analysis of Flowfields over Four-Engine DC-X Rockets

The objective of this study is to validate a computational methodology for the aerodynamic performance of an advanced conical launch vehicle configuration. The computational methodology is based on a three-dimensional, viscous flow, pressure-based computational fluid dynamics formulation. Both wind-tunnel and ascent flight-test data are used for validation. Emphasis is placed on multiple-engine power-on effects. Computational characterization of the base drag in the critical subsonic regime is the focus of the validation effort; until recently, almost no multiple-engine data existed for a conical launch vehicle configuration. Parametric studies using high-order difference schemes are performed for the cold-flow tests, whereas grid studies are conducted for the flight tests. The computed vehicle axial force coefficients, forebody, aftbody, and base surface pressures compare favorably with those of tests. The results demonstrate that with adequate grid density and proper distribution, a high-order difference scheme, finite rate afterburning kinetics to model the plume chemistry, and a suitable turbulence model to describe separated flows, plume/air mixing, and boundary layers, computational fluid dynamics is a tool that can be used to predict the low-speed aerodynamic performance for rocket design and operations.

Wang, Ten-See↗

NASA NTRS1998

Analysis of Linear Aerospike Plume Induced X-33 Base Heating Environment

Computational analysis is conducted to study the effect of an linear aerospike engine plume on the X-33 base-heating environment during ascent flight. To properly account for the freestream-body interaction and to allow for potential plume-induced flow-separation, the thermo-flowfield of the entire vehicle at several trajectory points is computed. A sequential grid-refinement technique is used in conjunction with solution-adaptive, patched, and embedded grid methods to limit the model to a manageable size. The computational methodology is based on a three-dimensional, finite-difference, viscous flow, chemically reacting, pressure-based computational fluid dynamics formulation, and a three-dimensional, finite-volume, spectral-line based weighted-sum-of-gray-gases absorption, computational radiation heat transfer formulation. The computed forebody and afterbody surface pressure coefficients and base pressure characteristic curves are compared with those of a cold-flow test. The predicted convective and radiative base-heat fluxes, the effect of base-bleed, and the potential of plume-induced flow separation are presented.

Wang, Ten-See↗

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 NTRS1999

Overview of Fluid Dynamics Activities at the Marshall Space Flight Center

Since its inception 40 years ago, Marshall Space Flight Center (MSFC) has had the need to maintain and advance state-of-the-art flow analysis and cold-flow testing capability to support its roles and missions. This overview discusses the recent organizational changes that have occurred at MSFC with emphasis on the resulting three groups that form the core of fluid dynamics expertise at MSFC: the Fluid Physics and Dynamics Group, the Applied Fluid Dynamics Analysis Group, and the Experimental Fluid Dynamics Group. Recently completed activities discussed include the analysis and flow testing in support of the Fastrac engine design, the X-33 vehicle design, and the X34 propulsion system design. Ongoing activities include support of the RLV vehicle design, Liquid Fly Back Booster aerodynamic configuration definition, and RLV focused technologies development. Other ongoing activities discussed are efforts sponsored by the Center Director's Discretionary Fund (CDDF) to develop an advanced incompressible flow code and to develop optimization techniques. Recently initiated programs and their anticipated required fluid dynamics support are discussed. Based on recent experiences and on the anticipated program needs, required analytical and experimental technique improvements are presented. Due to anticipated budgetary constraints, there is a strong need to leverage activities and to pursue teaming arrangements in order to advance the state-of-the-art and to adequately support concept development. Throughout this overview there is discussion of the lessons learned and of the capabilities demonstrated and established in support of the hardware development programs.

Garcia, Roberto↗

NASA NTRS1966

The Properties of Flowing Sheets Formed by Impingement of Liquid Jets on Curved Surfaces

An applied research program was conducted to determine the properties of flat sheets of propellants formed by directing jets of liquid tangentially against concave, cylindrical deflector surfaces. The dimensions and spatial orientation of the resultant sheets were found to depend only on the orifice diameter and deflector geometry for three propellant simulants of widely-varying physical properties, over the range of injection velocities encountered in liquid rocket engines. Correlating equations, suitable for use in injector design, are presented for free-sheet width and spreading angle. Distribution of mass and velocity across the free-flowing sheets is also reported. Conditions were discovered under which true sheets do not form, or at best malformed or pulsating sheets result. An envelope of geometrical constraints for deflector design is tentatively defined in order to avoid these undesirable operating regions. The results of the present cold-flow tests are compared with firing test data for impinging-sheet injectors and combustion performance is related to single sheet behavior.

Riebling, Robert W.↗

NASA NTRS1999

Flow Simulation of Solid Rocket Motors: Injection Induced Water-Flow Tests from Porous Media - 1

Prior to selecting a proper porous material for use in simulating the internal port flow of a solid rocket motor (SRM), in cold-flow testing, the flow emerging from porous materials is experimentally investigated. The injection-flow emerging from a porous matrix always exhibits a lumpy velocity profile that is spatially stable and affects the development of the longitudinal port flow. This flow instability, termed pseudoturbulence, is an inherent signature of the porous matrix and is found to generally increase with the wall porosity and with the injection flow rate. Visualization studies further show that the flow from porous walls made from shaving-type material (sintered stainless-steel) exhibits strong recirculation zones that are conspicuously absent in walls made from nodular or spherical material (sintered bronze). Detailed flow visualization observations and hot-film measurements are reported from tests of injection-flow and a coupled cross-flow from different porous wall materials. Based on the experimental data, discussion is provided on the choice of suitable material for SRM model testing while addressing the consequences and shortcomings from such a test.

Ramachandran, N.↗

NASA NTRS2000

Flow Simulation of Solid Rocket Motors: Sub-Scale Air Flow Simulation of Port Flows - 2

The injection-flow issuing from a porous medium in the cold-flow simulation of internal port flows in solid rocket motors is characterized by a spatial instability termed pseudoturbulence that produces a rather non-uniform (lumpy) injection-velocity profile. The objective of this study is to investigate the interaction between the injection- and the developing axial-flows. The findings show that this interaction generally weakens the lumpy injection profile and affects the subsequent development of the axial flow. The injection profile is found to depend on the material characteristics, and the ensuing pseudoturbulence is a function of the injection velocity, the axial position and the distance from the porous wall. The flow transition (from laminar to turbulent) of the axial-flow is accelerated in flows emerging from smaller pores primarily due to the higher pseudoturbulence produced by the smaller pores in comparison to that associated with larger pores. In flows with rather uniform injection-flow profiles (weak or no pseudoturbulence), the axial and transverse velocity components in the porous duct are found to satisfy the sine/cosine analytical solutions derived from inviscid assumptions. The transition results from the present study are compared with previous results from surveyed literature, and detailed flow development measurements are presented in terms of the blowing fraction, and characterizing Reynolds numbers.

Yeh, Y. P.↗