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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Vortex shedding from solid rocket propellant inhibitors

Vortex shedding frequency caused by the protrusion of inhibitors into the flow field of a solid rocket motor is investigated by experimental and mathematical models. The time dependent Navier-Stokes equations are solved using a finite difference technique assuming incompressible, two-dimensional flow under both laminar and turbulent flow conditions. For laminar flow, explicit solutions are obtained using a vorticity-transport equation in place of the Navier-Stokes equations. For turbulent flow, a two-equation (k-epsilon) model is used for turbulent modeling and the SIMPLE algorithm is employed as the computational scheme. Cold flow tests were conducted to confirm the basic flow structure and to determine the vortex shedding frequency under both laminar and turbulent flow conditions. The vortex shedding frequencies were determined using a stroboscope to measure the oscillating frequency of yarn tufts which were fastened to one inhibitor in the models. A hot-film anemometer established the velocity history behind the inhibitor. Good agreement between the theoretical results and measurements of the vortex shedding frequencies is demonstrated.

Shu, P. H.↗
NASA NTRS1989

Navier-Stokes analysis of solid propellant rocket motor internal flows

A multidimensional implicit Navier-Stokes analysis that uses numerical solution of the ensemble-averaged Navier-Stokes equations in a nonorthogonal, body-fitted, cylindrical coordinate system has been applied to the simulation of the steady mean flow in solid propellant rocket motor chambers. The calculation procedure incorporates a two-equation (k-epsilon) turbulence model and utilizes a consistently split, linearized block-implicit algorithm for numerical solution of the governing equations. The code was validated by comparing computed results with the experimental data obtained in cylindrical-port cold-flow tests. The agreement between the computed and experimentally measured mean axial velocities is excellent. The axial location of transition to turbulent flow predicted by the two-equation (k-epsilon) turbulence model used in the computations also agrees well with the experimental data. Computations performed to simulate the axisymmetric flowfield in the vicinity of the aft field joint in the Space Shuttle solid rocket motor using 14,725 grid points show the presence of a region of reversed axial flow near the downstream edge of the slot.

Sabnis, J. S.↗
NASA NTRS1991

Characterization of MOCVD fluid dynamics by laser velocimetry

Three-component laser velocimetry techniques are applied to characterize the fluid dynamics of an MOCVD reactor. These methods provide three-dimensional quantitative measurements of the gas velocities inside the reactor at typical growth temperatures. The effects of buoyancy-induced convection are examined by comparing data from the hot reactor with cold reactor results. It was found that thermal convection dominated the fluid dynamics of the reactor at growth temperatures. Cold reactor tests showed unstable flow patterns that were subject to fluidic switching effects.

Johnson, E. J.↗
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 and theoretical comparison of the Probe Thrust Vector Control concept

A concept that offers an alternate method for thrust vector control of liquid or solid propellant rockets is the use of a solid body or probe that is inserted on demand through the wall of the rocket nozzle. This Probe Thrust Vector Control (PTVC) concept is an alternative to that of a gimbaled nozzle or a Liquid Injection Thrust Vector control system. The viability of the PTVC concept can be assessed either experimentally and/or with the use of CFD. A purely experimental assessment is time consuming and expensive, whereas a CFD assessment is time- and cost-effective. Two key requirements of the concept are PTVC vectoring performance and active cooling requirements for the probe to maintain its thermal and structural integrity. The objective of the work reported here is presentation of experimental subscale cold flow tests and comparison of these tests with CFD predictions and the response time of the PTVC system.

Cavalleri, Robert↗
NASA NTRS1991

Transient response of a high-capacity heat pipe for Space Station Freedom

High-capacity heat pipe radiator panels have been proposed as the primary means of heat rejection for Space Station Freedom. In this system, the heat pipe would interface with the thermal bus condensers. Changes in system heat load can produce large temperature and heat load variations in individual heat pipes. Heat pipes could be required to start from an initially cold state, with heat loads temporarily exceeding their low-temperature transport capacity. The present research was motivated by the need for accurate prediction of such transient operating conditions. In this work, the cold startup of a 6.7-meter long high-capacity heat pipe is investigated experimentally and analytically. A transient thermohydraulic model of the heat pipe was developed which allows simulation of partially-primed operation. The results of cold startup tests using both constant temperature and constant heat flux evaporator boundary conditions are shown to be in good agreement with predicted transient response.

Ambrose, J. H.↗
NASA NTRS1991

Numerical analysis of ignition transients in solid rocket motors

A model to analyze the unstready multidimensional turbulent flow in the head-end, star slot section of a solid rocket motor during the ignition transient is developed. The present paper examines the fluid dynamic aspects of the starting transient. A finite-difference solution of the unsteady, compressible, full Navier-Stokes equations, together with a two-equation k-epsilon turbulence model, is obtained utilizing the MacCormack explicit, predictor-corrector technique. Computed results for the flow in the slot are compared with experimental data obtained in cold flow tests of a scaled model of the Space Shuttle SRM head-end, star grain section. The agreement between the calculated flowfield and existing experimental data is very good.

Ciucci, A.↗
NASA NTRS1991

Fluid management technology: Liquid slosh dynamics and control

Flight experiments were defined for the Cryogenic On-Orbit Liquid Depot Storage, Acquisition and Transfer Satellite (COLD-SAT) test bed satellite and the Shuttle middeck to help establish the influence of the gravitational environment on liquid slosh dynamics and control. Several analytical and experimental studies were also conducted to support the experiments and to help understand the anticipated results. Both FLOW-3D and NASA-VOF3D computer codes were utilized to simulate low Bond number, small amplitude sloshing, for which the motions are dominated by surface forces; it was found that neither code provided a satisfactory simulation. Thus, a new analysis of low Bond number sloshing was formulated, using an integral minimization technique that will allow the assumptions made about surface physics phenomena to be modified easily when better knowledge becomes available from flight experiments. Several examples were computed by the innovative use of a finite-element structural code. An existing spherical-pendulum analogy of nonlinear, rotary sloshing was also modified for easier use and extended to low-gravity conditions. Laboratory experiments were conducted to determine the requirements for liquid-vapor interface sensors as a method of resolving liquid surface motions in flight experiments. The feasibility of measuring the small slosh forces anticipated in flight experiments was also investigated.

Dodge, Franklin T.↗
NASA NTRS1992

Experimental studies of hypersonic shock-wave boundary-layer interactions

Two classes of shock-wave boundary-layer interactions were studied experimentally in a shock tunnel in which a low Reynolds number, turbulent flow at Mach 8 was developed on a cold, flat test surface. The two classes of interactions were: (1) a swept interaction generated by a wedge ('fin') mounted perpendicularly on the flat plate; and (2) a two-dimensional, unseparated interaction induced by a shock impinging near an expansion corner. The swept interaction, with wedge angles of 5-20 degrees, was separated and there was also indication that the strongest interactions prossessed secondary separation zones. The interaction spread out extensively from the inviscid shock location although no indication of quasi-conical symmetry was evident. The surface pressure from the upstream influence to the inviscid shock was relatively low compared to the inviscid downstream value but it rose rapidly past the inviscid shock location. However, the surface pressure did not reach the downstream inviscid value and reasons were proposed for this anomalous behavior compared to strongly separated, supersonic interactions. The second class of interactions involved weak shocks impinging near small expansion corners. As a prelude to studying this interaction, a hypersonic similarity parameter was identified for the pure, expansion corner flow. The expansion corner severely damped out surface pressure fluctuations. When a shock impinged upstream of the corner, no significant changes to the surface pressure were found as compared to the case when the shock impinged on a flat plate. But, when the shock impinged downstream of the corner, a close coupling existed between the two wave systems, unlike the supersonic case. This close coupling modified the upstream influence. Regardless of whether the shock impinged ahead or behind the corner, the downstream region was affected by the close coupling between the shock and the expansion. Not only was the mean pressure distribution modified but the unsteadiness in the surface pressure was reduced compared to the flat-plate case.

Lu, Frank K.↗
NASA NTRS1992

Results of an experimental investigation of the flow field in the head-end star slot section of a Solid Rocket Motor

Cold flow tests of a one-tenth scale model based on the geometry of the Space Shuttle Solid Rocket Motor, SRM, head-end section and its single-port igniter, along with two models of a four port igniter, have been conducted. The tests were done to establish quantitative as well as qualitative data on the behavior of the flow inside the star slot during the early part of the ignition transient. This paper presents the data obtained from these tests and discusses the implications of the data with respect to ignition transients of solid rocket motors which have head-end star grains.

Ciucci, A.↗
NASA NTRS1992

Thermostructural applications of heat pipes for cooling leading edges of high-speed aerospace vehicles

Heat pipes have been considered for use on wing leading edge for over 20 years. Early concepts envisioned metal heat pipes cooling a metallic leading edge. Several superalloy/sodium heat pipes were fabricated and successfully tested for wing leading edge cooling. Results of radiant heat and aerothermal testing indicate the feasibility of using heat pipes to cool the stagnation region of shuttle-type space transportation systems. The test model withstood a total seven radiant heating tests, eight aerothermal tests, and twenty-seven supplemental radiant heating tests. Cold-wall heating rates ranged from 21 to 57 Btu/sq ft-s and maximum operating temperatures ranged from 1090 to 1520 F. Follow-on studies investigated the application of heat pipes to cool the stagnation regions of single-stage-to-orbit and advanced shuttle vehicles. Results of those studies indicate that a 'D-shaped' structural design can reduce the mass of the heat-pipe concept by over 44 percent compared to a circular heat-pipe geometry. Simple analytical models for heat-pipe startup from the frozen state (working fluid initially frozen) were adequate to approximate transient, startup, and steady-state heat-pipe performance. Improvement in analysis methods has resulted in the development of a finite-element analysis technique to predict heat-pipe startup from the frozen state. However, current requirements of light-weight design and reliability suggest that metallic heat pipes embedded in a refractory composite material should be used. This concept is the concept presently being evaluated for NASP. A refractory-composite/heat-pipe-cooled wing leading edge is currently being considered for the National Aero-Space Plane (NASP). This concept uses high-temperature refractory-metal/lithium heat pipes embedded within a refractory-composite structure and is significantly lighter than an actively cooled wing leading edge because it eliminates the need for active cooling during ascent and descent. Since the NASP vehicle uses cryogenic hydrogen to cool structural components and then burns this fuel in the combustor, hydrogen necessary for descent cooling only, when the vehicle is unpowered, is considered to be a weight penalty. Details of the design of the refractory-composite/heat-pipe-cooled wing leading edge are currently being investigated. Issues such as thermal contact resistance and thermal stress are also being investigated.

Camarda, Charles J.↗
NASA NTRS1992

Experimental investigation of a simulated LOX injector flow field and other nonintrusive measurement efforts

Efforts to improve the characteristics of fuel-oxidizer mixing in liquid rocket combustors have lead to a swirl element design for a liquid oxygen injector which is being considered for use on the STME. For the design which is the subject of this investigation, the oxygen enters the injector element perpendicular to the injector axis and nearly tangent to the circular injector wall. This swirl element is at one end of a tube and the injector exit is at the other. This geometric configuration creates a plume in the shape of a conical sheet. This sheet is either primarily contiguous liquid or droplets depending on the pressure drop in the injector and the distance from the injector exit. Probe-based devices such as two-dimensional grid patternators have been used to investigate simulated LOX injector flow fields (Hulka). The primary work described herein is an effort to use optical techniques to investigate the plume of a swirl injector element. For this investigation, a high pressure (500 psig) cold flow test facility was constructed. Water was used as the LOX simulate and air pressure was used to drive the injector flow field. Laser-induced fluorescence (LIF) from dye seeded into the water was used to obtain quantitative measurements of the time-averaged water concentration distribution in the plume. Scattered laser light and LIF were used for time averaged plume visualization and scattered light from a strobe with a 1 microsecond pulse was used for time-resolved plume visualization. During the Summer Faculty Fellowship for which this report was developed, an additional effort, unrelated to the swirl injector investigation, was made to resolve fluctuations in the combustion product composition in the exhaust of a hybrid rocket motor. A brief description of this effort is included herein.

Hartfield, Roy J., Jr.↗
NASA NTRS1992

Investigation of rotor blade roughness effects on turbine performance

The cold air test program was completed on the SSME (Space Shuttle Main Engine) HPFTP (High Pressure Fuel Turbopump) turbine with production nozzle vane rings and polished coated rotor blades with a smooth surface finish of 30 microinch (0.76 micrometer) RMS (Root Mean Square). The smooth blades were polished by an abrasive flow machining process. The test results were compared with the air test results from production rough coated rotor blades with a surface finish of up to 400 microinch (10.16 micrometer) RMS. Turbine efficiency was higher for the smooth blades over the entire range tested. Efficiency increased 2.1 percentage points at the SSME 104 percent RPL (Rated Power Level) condition. This efficiency improvement could reduce the SSME HPFTP turbine inlet temperature by 57 degrees Rankine (32 degrees Kelvin) increasing turbine durability. The turbine flow parameter increased and the mid-span outlet swirl angle became more axial with the smooth rotor blades.

Boynton, J. L.↗
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 NTRS1992

Design and evaluation of single and dual flow thrust vector nozzles with post exit vanes

This Thrust Vectored Research project required that a 1/24 scale model of the F/A-18 High Alpha Research Vehicle, (HARV), propulsion system be constructed on the university campus. This propulsion system was designed for cold flow testing on a multicomponent test rig. Forces and moments were measured to study nozzle performance parameters. The flow visualization technique of color Schlieren photography was performed to investigate the flow phenomena at the nozzle exit. The flow interactions that were identified consisted of vane nozzleing between the outer and lower vanes and vane tip interference. The thrust vectoring system consisted of three asymmetrically spaced vanes installed circumferentially on a maximum afterburner nozzle. The performance of the nozzle was investigated with the outer and lower vanes equally deflected, (-10 deg is less than delta(sub v) is less than 25 deg), and with the upper vane fully retracted, (delta(sub v) equals -10 deg). The nozzle pressure ratio ranged from 4 to 6. The results indicated that a vane nozzleing effect developed at nozzle pressure ratios of 4 and 6 when the outer and lower vanes were deflected far enough into the flow field such that the increase in vane area accelerated the flow past the vanes causing distorted shock waves. This accelerated flow was a result of a pressure differential existing between the inside surface of the vane and the ambient pressure. The stagnation pressure that developed along the inside surface of the vane accelerated the flow past the vanes causing it to equalize with ambient pressure, thus providing distorted shock waves. A tip interference was present at the trailing edge of the upper vane as a result of low nozzle pressure, NPR 4, with high vane deflection, delta(sub v) equals 25 degrees, and also with a high nozzle pressure, NPR 6, and low vane deflections, delta(sub v) equals 15 degrees.

Carpenter, Thomas W.↗
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 NTRS1993

Dual-bell altitude compensating nozzles

The primary objective of this cold flow test effort was to assess the performance characteristics of dual bell nozzles and to obtain preliminary design criteria by testing a number of configurations. Characteristics of interest included low altitude performance, high altitude performance, and the flow transition process. In combination with this performance data, other factors such as cost, weight, fabricability, and vehicle related issues could then be traded to establish the feasibility of the concept.

Horn, M.↗
NASA NTRS1991

Orbit transfer rocket engine technology program enhanced heat transfer combustor technology

In order to increase the performance of a high performance, advanced expander-cycle engine combustor, higher chamber pressures are required. In order to increase chamber pressure, more heat energy is required to be transferred to the combustor coolant circuit fluid which drives the turbomachinery. This requirement was fulfilled by increasing the area exposed to the hot-gas by using combustor ribs. A previous technology task conducted 2-d hot air and cold flow tests to determine an optimum rib height and configuration. In task C.5 a combustor calorimeter was fabricated with the optimum rib configuration, 0.040 in. high ribs, in order to determine their enhancing capability. A secondary objective was to determine the effects of mixture ratio changers on the enhancement during hot-fire testing. The program used the Rocketdyne Integrated Component Evaluator (ICE) reconfigured into a thrust chamber only mode. The test results were extrapolated to give a projected enhancement from the ribs for a 16 in. long cylindrical combustor at 15 Klb nominal thrust level. The hot-gas wall ribs resulted in a 58 percent increase in heat transfer. When projected to a full size 15K combustor, it becomes a 46 percent increase. The results of those tests, a comparison with previous 2-d results, the effects of mixture ratio and combustion gas flow on the ribs and the potential ramifications for expander cycle combustors are detailed.

Brown, William S.↗