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NASA NTRS1999

Control of a Thrust Alignment Table for Modeling the Coning Dynamics of a Spinning Spacecraft with a Follower Force

This document presents a system controlling the motion of a spherical air bearing used in the modeling of spacecraft dynamics and controls in a laboratory environment. The system is part of the Spinning Rocket Simulator (SRS), used to simulate the coning of spacecraft during a thrusting stage. The reaction force at the spherical air bearing supporting the spacecraft model must coincide with the thrust axis of the model for proper simulation. Therefore, the bearing is translated in a circular path to introduce a centrifugal force. This horizontal force along with the gravitational reaction force at the bearing combines to simulate the direction of the spacecraft's thrust force. The control system receives attitude information from the spacecraft model via a laser beam embedded in the model that impinges on a photosensitive array. The non-linear system is controlled using high-speed lookup tables and digital techniques. A vector-controlled motor and a stepper motor are given the necessary signals to accurately control the turntable and platform supporting the air bearing. Preliminary performance data is presented. Mechanical elements of the table and platform are described in detail. A wireless (RF) data path for all devices on the spacecraft model to an off-table command computer is also described.

Halsmer, Dominic↗

NASA NTRS1998

The Applicability of Jet-Shear-Layer Mixing and Effervescent Atomization for Low-NO(x) Combustors

An investigation has been conducted to develop appropriate technologies for a low-NO(x), liquid-fueled combustor. The combustor incorporates an effervescent atomizer used to inject fuel into a premixing duct. Only a fraction of the combustion air is used in the premixing process. This fuel-rich mixture is introduced into the remaining combustion air by a rapid jet-shear-layer mixing process involving radial fuel-air jets impinging on axial air jets in the primary combustion zone. Computational modeling was used as a tool to facilitate a parametric analysis appropriate to the design of an optimum low-NO(x) combustor. A number of combustor configurations were studied to assess the key combustor technologies and to validate the three-dimensional modeling code. The results from the experimental testing and computational analysis indicate a low-NO(x) potential for the jet-shear-layer combustor. Key features found to affect NOx emissions are the primary combustion zone fuel-air ratio, the number of axial and radial jets, the aspect ratio and radial location of the axial air jets, and the radial jet inlet hole diameter. Each of these key parameters exhibits a low-NO(x) point from which an optimized combustor was developed Also demonstrated was the feasibility of utilizing an effervescent atomizer for combustor application. Further developments in the jet-shear-layer mixing scheme and effervescent atomizer design promise even lower NO(x) with high combustion efficiency.

Colantonio, R. O.↗

NASA NTRS1999

Blade Heat Transfer Measurements and Prediction in a Transonic Turbine Cascade

Detailed heat transfer measurements and predictions are given for a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. Data were obtained for inlet Reynolds numbers of 0.5 and 1.0 x 10(exp 6), for isentropic exit Mach numbers of 1.0 and 1.3, and for inlet turbulence intensities of 0.25% and 7.0%. Measurements were made in a linear cascade having a highly three-dimensional flow field resulting from thick inlet boundary layers. The purpose of the work is to provide benchmark quality data for three-dimensional CFD code and model verification. Data were obtained by a steady-state technique using a heated, isothermal blade. Heat fluxes were determined from a calibrated resistance layer in conjunction with a surface temperature measured by calibrated liquid crystals. The results show the effects of strong secondary vortical flows, laminar-to-turbulent transition, shock impingement, and increased inlet turbulence on the surface heat transfer.

Giel, P. W.↗

NASA NTRS1999

Optical Imaging of Flow Pattern and Phantom

Time-resolved optical imaging technique has been used to image the spatial distribution of small droplets and jet sprays in a highly scattering environment. The snake and ballistic components of the transmitted pulse are less scattered, and contain direct information about the sample to facilitate image formation as opposed to the diffusive components which are due to multiple collisions as a light pulse propagates through a scattering medium. In a time-gated imaging scheme, these early-arriving, image-bearing components of the incident pulse are selected by opening a gate for an ultrashort period of time and a shadowgram image is detected. Using a single shot cooled CCD camera system, the formation of water droplets is monitored as a function of time. Picosecond time-gated image of drop in scattering cells, spray droplets as a function of let speed and gas pressure, and model calcification samples consisted of calcium carbonate particles of irregular shapes ranging in size from 0. 1 to 1.5 mm affixed to a microscope slide have been measured. Formation produced by an impinging jet will be further monitored using a CCD with 1 kHz framing illuminated with pulsed light. The desired image resolution of the fuel droplets is on the 20 pm scale using early light through a highly scattering medium. A 10(exp -6)m displacement from a jet spray with a flow speed of 100 m/sec introduced by the ns grating pulse used in the imaging is negligible. Early ballistic/snake light imaging offers nondestructive and noninvasive method to observe the spatial distribution of hidden objects inside a highly scattering environment for space, biomedical, and materials applications. In this paper, the techniques we will present are time-resolved K-F transillumination imaging and time-gated scattered light imaging. With a large dynamic range and high resolution, time-gated early light imaging has the potential for improving rocket/aircraft design by determining jets shape and particle sizes. Refinements to these techniques may enable drop size measurements in the highly scattering, optically dense region of multi-element rocket injectors. These types of measurements should greatly enhance the design of stable, and higher performing rocket engines.

Galland, Pierre A.↗

NASA NTRS1999

Pressure-Sensitive Paint Measurements on Surfaces with Non-Uniform Temperature

Pressure-sensitive paint (PSP) has become a useful tool to augment conventional pressure taps in measuring the surface pressure distribution of aerodynamic components in wind tunnel testing. While the PSP offers the advantage of a non-intrusive global mapping of the surface pressure, one prominent drawback to the accuracy of this technique is the inherent temperature sensitivity of the coating's luminescent intensity. A typical aerodynamic surface PSP test has relied on the coated surface to be both spatially and temporally isothermal, along with conventional instrumentation for an in situ calibration to generate the highest accuracy pressure mappings. In some tests however, spatial and temporal thermal gradients are generated by the nature of the test as in a blowing jet impinging on a surface. In these cases, the temperature variations on the painted surface must be accounted for in order to yield high accuracy and reliable data. A new temperature correction technique was developed at NASA Lewis to collapse a "family" of PSP calibration curves to a single intensity ratio versus pressure curve. This correction allows a streamlined procedure to be followed whether or not temperature information is used in the data reduction of the PSP. This paper explores the use of conventional instrumentation such as thermocouples and pressure taps along with temperature-sensitive paint (TSP) to correct for the thermal gradients that exist in aeropropulsion PSP tests. Temperature corrected PSP measurements for both a supersonic mixer ejector and jet cavity interaction tests are presented.

Bencic, Timothy J.↗

NASA NTRS2000

Raman Gas Species Measurements in Hydrocarbon-Fueled Rocket Engine Injector Flows

Propellent injector development at MSFC (Marshall Space Flight Center) includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellent mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

Wehrmeyer, Joseph A.↗

NASA NTRS1999

Liquid Methane/Liquid Oxygen Injectors for Potential Future Mars Ascent Engines

Preliminary mission studies for human exploration of Mars have been performed at Marshall Space Flight Center (MSFC). These studies indicate that for chemical rockets only a cryogenic propulsion system would provide high enough performance to be considered for a Mars ascent vehicle. Although the mission is possible with Earth-supplied propellants for this vehicle, utilization of in-situ propellants is highly attractive. This option would significantly reduce the overall mass of launch vehicles. Consequently, the cost of the mission would be greatly reduced because the number and size of the Earth launch vehicle(s) needed for the mission would decrease. NASA/Johnson Space Center has initiated several concept studies of in-situ propellant production plants. Liquid oxygen (LOX) is the primary candidate for an in-situ oxidizer. In-situ fuel candidates include methane (CH4), ethylene (C2H4), and methanol (CH3OH). MSFC initiated a technology development program for a cryogenic propulsion system for the Mars human exploration mission in 1998. One part of this technology program is the effort described here: an evaluation of propellant injection concepts for a LOX/liquid methane Mars Ascent Engine (MAE) with an emphasis on light-weight, high efficiency, reliability, and thermal compatibility. In addition to the main objective, hot-fire tests of the subject injectors will be used to test other key technologies including light-weight combustion chamber materials and advanced ignition concepts. This paper will address the results of the liquid methane/LOX injector study conducted at MSFC. A total of four impinging injector configurations were tested under combustion conditions in a modular combustor test article (MCTA), equipped with optically accessible windows. A series of forty hot-fire tests, which covered a wide range of engine operating conditions with the chamber pressure varied from 320 to 510 and the mixture ratio from 1.5 to 3.5, were performed. The test matrix also included a variation in the combustion chamber length for the purpose of investigating its effects on the combustion performance and stability.

Trinh, Huu Phuoc↗

NASA NTRS1999

Growth of Solid Solution Single Crystals

The solidification of a solid solution semiconductor, having a wide separation between liquidus and serious has been extensively studied in ground based, high magnetic field and Spacelab experiments. Two alloys of mercury cadmium telluride have been studied; mercury cadmium telluride with 80.0 mole percent of HgTe and 84.8 mole percent respectively. These alloys are extremely difficult to grow by directional solidification on earth due to high solutal and thermal density differences that give rise to fluid flow and consequent loss of interface shape and composition. Diffusion controlled growth is therefore impossible to achieve in conventional directional solidification. The ground based experiments consisted of growing crystals in several different configurations of heat pipe furnaces, NASA's Advanced Automated Directional Solidification Furnace (AADSF), and a similar furnace incorporated in a superconducting magnet capable of operating at up to 5T. The first microgravity experiment took place during the flight of STS-62 in March 1994, with the AADSF installed on the second United States Microgravity Payload (USMP-2). The alloy was solidified at 3/4 inch per day over a 9 day period, and for the first time a detailed evaluation was performed of residual acceleration effects. The second flight experiment took place in the fourth United States Microgravity Payload Mission (USMP-4) in November 1997. Due to contamination of the furnace system by a previously processed sample, the sample was not received until May 1998, and the preliminary analysis shows that the conditions prevailing during the experiment were quite different from the requirements requested prior to the mission. Early results are indicating that the sample may not accomplish the desired objectives. As with the USMP-2 mission, the results of the ground based experiments were compared with the crystal grown in orbit under microgravity conditions. On the earth, it has been demonstrated that the application of the magnetic field leads to a significant reduction in fluid flow, with improved homogeneity of composition. The field strength required to suppress flow increases with diameter of the material. The 8 mm diameter sample used here was less than the upper diameter limit for a ST magnet. The configuration for USMP-4 was changed so that the material was seeded and other processing techniques were also modified. It was decided to examine the effects of a strong magnetic field under the modified configuration and parameters. A further change from USMP-2 was that a different composition of material was grown, namely with 0.152 mole fraction of cadmium telluride rather than the 0.200 of the USMP-2 experiment. The objective was to grow highly homogeneous, low defect density material of a composition at which the conduction band and the valence band of the material impinge against each other. As indicated, the furnace was contaminated during the mission. As a result of solid debris remaining in the furnace bore, the cartridge in this experiment, denoted as SL1-417, was significantly bent during the insertion phase. During translation the cartridge scraped against the plate which isolates the hot and cold zones of the furnace. Thermocouples indicated that a thermal assymetry resulted. The scraping in the slow translation or crystal growth part of the processing was not smooth and it is probable that the jitter was sufficient to give rise to convection in the melt. Early measurements of composition from the surface of the sample have shown that the composition varies in an oscillatory manner.

Lehoczky, Sandor L.↗

NASA NTRS1999

The Dipper Satellite: A Medium-Class Explorer Mission to the Threshold of Space

The Dipper satellite will carry out an unprecedented, systematic, and focused in-situ exploration of the Earth's lower ionosphere and thermosphere below 200 km that will produce a pivotal base of knowledge that will significantly advance our understanding of knowledge that will significantly advance our understanding of how our near-space environment works. The satellite will carry comprehensive in-situ probes to measure vector electric and magnetic fields, plasma density and temperature, ion velocities, ion and neutral composition and winds, energetic particles including suprathermal electrons, gravity waves, and lightning bursts. The satellite will include a propulsion system and tapered body that will provide over 10,000 excursions to altitudes below 200 km with over 3000 dips to altitudes below 150 km. With this instrument complement, spacecraft, and orbit, the Dipper mission will gather the necessary combined electrodynamics and neutral dynamics measurements to provide an understanding of the Earth's critical boundary region where the ionized gases of space and the neutral gases of the atmosphere are coupled, and where impinging forces and momentum are deposited from the magnetosphere above and from the troposphere, stratosphere, and mesosphere below. In exploring those physical processes in the lower ionosphere which can only be measured in-situ, the Dipper mission addresses four main science objectives. The Dipper will: 1) reveal how ion-neutral coupling creates a global system of dynamo electric fields and currents; 2) provide first-hand understanding of how magnetospheric currents close in the ionosphere and reveal the effects on the upper atmosphere of magnetospheric energy and momentum deposition; 3) discover the degree of upwards coupling and energy deposition due to thunderstorm electric fields and determine their significance; 4) determine the dynamics and composition of the Earth's lower thermosphere, including its response to gravity, tidal, and planetary waves on a range of spatial scales. A proposal to design, build, operate, and analyze data from instruments on the Dipper spacecraft within the schedule and budget constraints of NASA's MIDEX program was submitted to NASA in 1998. This presentation summarizes the main features of the mission.

Pfaff, R. F., Jr.↗

NASA NTRS1999

Liquid Methane/Oxygen Injector Study for Mars Ascent Engines

As a part of the advancing technology of the cryogenic propulsion system for the Mars exploration mission, this effort aims at evaluating propellant injection concepts for liquid methane/liquid oxygen (LOX) rocket engines. Split-triplet and unlike impinging injectors were selected for this study. A total of four injector configurations were tested under combustion conditions in a modular combustor test article (MCTA), equipped with optically accessible windows, at MSFC. A series of forty hot-fire tests, which covered a wide range of engine operating conditions with the chamber pressure ranging from 320 to 510 and the mixture ratio from 1.5 to 3.5, were conducted. The test matrix also included a variation in the combustion chamber length for the purpose of investigating its effects on the combustion performance and stability. Initial assessments of the test results showed that the injectors provided stable combustion and there were no injector face overheating problems under all operating conditions. The Raman scattering signal measurement method was successfully demonstrated for the hydrocarbon/oxygen reactive flow field. The near-injector face flow field was visually observed through the use of an infrared camera. Chamber wall temperature, high frequency chamber pressure, and average throat section heat flux were also recorded throughout the test series. Assessments of the injector performance are underway.

Trinh, Huu Phuoc↗

NASA NTRS1999

Feasibility Assessment of Thermal Barriers for RSRM Nozzle Joint Locations

Solid rockets, including the Space Shuttle solid rocket motor, are generally manufactured in large segments which are then shipped to their final destination where they are assembled. These large segments are sealed with a system of primary and secondary 0-rings to contain combustion gases inside the rocket which are at pressures of up to 900 psi and temperatures of up to 5500 F. The seals are protected from hot combustion gases by thick layers of phenolic insulation and by joint-filling compounds between these layers. Recently, though, routine inspections of nozzle-to-case joints in the Shuttle solid rocket motors during disassembly revealed erosion of the primary O-rings. Jets of hot gas leaked through gaps in the joint-filling compound between the layers of insulation and impinged on the O-rings. This is not supposed to take place, so NASA and Thiokol, the manufacturer of the rockets, initiated an investigation and found that design improvements could be made in this joint. One such improvement would involve using NASA Lewis braided thermal barriers as another level of protection for the O-ring seals against the hot combustion gases.

Steinetz, Bruce M.↗

NASA NTRS1990

The Importance of the Measurement of the Unsteady Wake of the Aeroassist Flight Experiment Vehicle: A White Paper

The High Frequency Radiometer (HFR) is the only instrument on the Aeroassist Flight Experiment (AFE) with sufficient temporal resolution to discern the frequency of unsteady wake oscillations. Determining both the frequency and amplitude of wake unsteadiness during AFE atmospheric entry is essential for reliably predicting the geometry and motion of the wake of future Aeroassisted Space Transfer Vehicles (ASTV). These parameters directly affect the location and size of the pay- load and the weight of the required afterbody heat protection. The purpose of the AFE is to validate the technologies required for the design of ASTVs, which will be used to exploit Earth-lunar space. This validation will be conducted at a combination of vehicle size, altitude, and velocity not obtainable in ground-based facilities. The AFE will provide the experimental flight data needed to improve our understanding of hypersonic-wake physics and to validate computational predictions of the aerodynamic and heating loads, including afterbody radiative heating loads, on an ASTV. Reliable prediction of ASTV wake flows will ensure that payloads are located within the shear-layer envelope and will determine the amount of thermal protection the payloads require. Specifically, understanding the temporal nature of the wake unsteadiness is important for two reasons. Most importantly, analysis of ground-based experiments suggests that wake unsteadiness results in a variation of as much as +/- 5 deg in the shear-flow turning angle. This angle must be reliably predicted to avoid shear-layer impingement on the vehicle afterbody, which would result in heating rates of about 10 W/sq cm, of the same order as on the forebody stagnation point. Secondly, the energy associated with wake unsteadiness will reduce the static enthalpy of the wake fluid and cause an error of as much as 30% in the amount of predicted wake radiative heating. Therefore, the HFR flight data, which will quantify the frequency and amplitude of the wake unsteadiness, are required for the verification of computational models of ASTV flowfields. Measurement of wake unsteadiness has been endorsed by the Peer Science Steering Group and the AFE Computational Fluid Dynamics (CFD) Working Group. The Peer Science Steering Group has stressed that a reliable measurement of the wake unsteadiness is fundamental because incorporating unsteadiness represents a substantial challenge to the CFD community, and a reliable measurement will raise confidence in the computer simulation. The AFE CFD Working Group has stated that 44 any data which would prove the existence and effects of unsteady flow would be extremely valuable.

Strawa, A. W.↗

NASA NTRS1999

Titanium Optics for Ion Thrusters

Ion thruster total impulse capability is limited, in part, by accelerator grid sputter erosion. A development effort was initiated to identify a material with a lower accelerator grid volumetric sputter erosion rate than molybdenum, but that could utilize the present NSTAR thruster grid design and fabrication techniques to keep development costs low, and perform as well as molybdenum optics. After comparing the sputter erosion rates of several atomic materials to that of molybdenum at accelerator voltages, titanium was found to offer a 45% reduction in volumetric erosion rates. To ensure that screen grid sputter erosion rates are not higher at discharge chamber potentials, titanium and molybdenum sputter erosion rates were measured at these potentials. Preliminary results showed only a slightly higher volumetric erosion rate for titanium, so that screen grid erosion is insignificant. A number of material, thermal, and mechanical properties were also examined to identify any fabrication, launch environment, and thruster operation issues. Several titanium grid sets were successfully fabricated. A titanium grid set was mounted onto an NSTAR 30 cm engineering model ion thruster and tested to determine optics performance. The titanium optics operated successfully over the entire NSTAR power range of 0.5 to 2.3 kW. Differences in impingement-limited perveances and electron backstreaming limits were found to be due to a larger cold gap for the titanium optics. Discharge losses for titanium grids were lower than those for molybdenum, likely due to a slightly larger titanium screen grid open area fraction. Radial distributions of beam current density with titanium optics were very similar to those with molybdenum optics at all power levels. Temporal electron backstreaming limit measurements showed that titanium optics achieved thermal equilibrium faster than molybdenum optics.

Soulas, George C.↗

NASA NTRS2000

Applying Contamination Modelling to Spacecraft Propulsion Systems Designs and Operations

Molecular and particulate contaminants generated from the operations of a propulsion system may impinge on spacecraft critical surfaces. Plume depositions or clouds may hinder the spacecraft and instruments from performing normal operations. Firing thrusters will generate both molecular and particulate contaminants. How to minimize the contamination impact from the plume becomes very critical for a successful mission. The resulting effect from either molecular or particulate contamination of the thruster firing is very distinct. This paper will discuss the interconnection between the functions of spacecraft contamination modeling and propulsion system implementation. The paper will address an innovative contamination engineering approach implemented from the spacecraft concept design, manufacturing, integration and test (I&T), launch, to on- orbit operations. This paper will also summarize the implementation on several successful missions. Despite other contamination sources, only molecular contamination will be considered here.

Chen, Philip T.↗

NASA NTRS1999

CFD Predictions of Sonic-Boom Characteristics for Unmodified and Modified SR-71 Configurations

Shaped sonic-boom signatures refer to signatures that look something other than the typical N-waves. Shaped sonic-boom signatures such as "flat-top," "ramp-type," or "hybrid-type" waveforms have been shown to reduce the subjective loudness without requiring reductions in overpressure peaks. The shaping of sonic-boom signatures requires increasing the shock rise time and changes in frequency spectra. So far, a flat-top waveform was shown to be achievable in wind tunnels; however, the influence of long propagation distance and real atmosphere on shaped signatures should be addressed using flight tests. Two different approaches have been proposed for sonic-boom minimization flight tests. The first approach, proposed by Eagle Aerospace, is for a flight test using a modified BQM-34 "FIREBEE" remotely piloted vehicle. The 30-foot long FIREBEE has a steady state flight condition at the Mach number and altitude of interest, and it can be recovered by helicopter from the water. As an alternative approach, a modified SR-71 vehicle has been proposed by the McDonnell Douglas Corporation. Benefits of the SR-71 include its variable geometry supersonic inlets, small cockpit bulge, higher Mach number capabilities, slender design, and longer length (105 foot). The present investigation addresses the sonic-boom analysis for the second vehicle.The objective of the current investigation is to assess the feasibility of a modified SR-71 configuration, with McDonnell Douglas-designed fuselage modifications, intended to produce shaped sonic-boom signatures on the ground. The present study describes the use of a higher-order computational fluid dynamics (CFD) method to predict the sonic-boom characteristics for both unmodified and modified SR-71 configurations. An Euler unstructured grid methodology is used to predict the near-field, three-dimensional pressure patterns generated by both SR-71 models. The computed near-field pressure signatures are extrapolated to specified distances below the aircraft down to impingement on the ground using the code MDBOOM. Comparisons of the near-field pressure signatures with available flight-test data are presented in the current paper.

Fouladi, Kamran↗

NASA NTRS1999

A Design Tool for Liquid Rocket Engine Injectors

A practical design tool for the analysis of flowfields near the injector face has been developed and used to analyze the Fastrac engine. The objective was to produce a computational design tool which was detailed enough to predict the interactive effects of injector element impingement angles and points and the momenta of the individual orifice flows. To obtain a model which could be used to simulate a significant number of individual orifices, a homogeneous computational fluid dynamics model was developed. To describe liquid and vapor sub- and super-critical flows, the model included thermal and caloric equations of state which were valid over a wide range of pressures and temperatures. A homogeneous model was constructed such that the local state of the flow was determined directly, i.e. the quality of the flow was calculated. Such a model does not identify drops or their distribution, but it does allow the flow along the injector face and into the acoustic cavity to be predicted. It also allows the film coolant flow to be accurately described. The initial evaluation of the injector code was made by simulating cold flow from an unlike injector element and from a like-on-like overlapping fan (LOL) injector element. The predicted mass flux distributions of these injector elements compared well to cold flow test results. These are the same cold flow tests which serve as the data base for the JANNAF performance prediction codes. The flux distributions 1 inch downstream of the injector face are very similar; the differences were somewhat larger at further distances from the faceplate. Since the cold flow testing did not achieve good mass balances when integrations across the entire fan were made, the CFD simulation appears to be reasonable alternative to future cold flow testing. To simulate the Fastrac, an RP-1/LOX combustion model must be chosen. This submodel must be relatively simple to accomplish three-dimensional, multiphase flow simulations. Single RP-1 pyrolysis and partial oxidation steps were chosen and the combustion was completed with the wet CO mechanism. Soot was also formed with a single global reaction. To validate the combustion submodel, global data from gas generator tests and from subscale motor test were used to predict qualitatively correct mean molecular weights, temperature, and soot levels. Because such tests do not provide general kinetics rates, the methodology is not necessarily appropriate for other than rocket type flows conditions. Soot predictions were made so that radiation heating to the motor walls can be made. These initial studies of the Fastrac were for a small region close to the injector face and chamber wall which included a segment of the acoustic cavity. The region analyzed includes 11 individual orifice holes to represent the LOL elements and the H2 film coolant holes. Typical results of this simulation are shown in Figure 1. At this point the only available test data to verify the predictions are temperatures measured in the acoustic cavity. These temperatures are in reasonable agreement at about 2000R (1111 K). Future work is expected to include improving the computational efficiency or the CFD model and/or using more computer capacity than the single Pentium PC with which these simulations were made.

Farmer, Richard C.↗

NASA NTRS2000

Flow Environment Study Near the Empennage of a 15-Percent Scale Helicopter Model

Development of advanced rotorcraft configurations has highlighted a need for high-quality experimental data to support the development of flexible and accurate analytical design tools. To provide this type of data, a test program was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to measure the flow near the empennage of a 15-percent scale powered helicopter model with an operating tail fan. Three-component velocity profiles were measured with laser velocimetry (LV) one chord forward of the horizontal tail for four advance ratios to evaluate the effect of the rotor wake impingement on the horizontal tail angle of attack. These velocity data indicate the horizontal tail can experience unsteady angle of attack variations of over 30 degrees due to the rotor wake influence. The horizontal tail is most affected by the rotor wake above advance ratios of 0.10. Velocity measurements of the flow on the inlet side of the tail fan were made for a low-speed flight condition using conventional LV techniques. The velocity data show an accelerated flow near the tail fan duct, and vorticity calculations track the passage of main rotor wake vortices through the measurement plane.

Gorton, Susan Althoff↗

NASA NTRS2000

Numerical Simulation of Non-Rotating and Rotating Coolant Channel Flow Fields

Future generations of ultra high bypass-ratio jet engines will require far higher pressure ratios and operating temperatures than those of current engines. For the foreseeable future, engine materials will not be able to withstand the high temperatures without some form of cooling. In particular the turbine blades, which are under high thermal as well as mechanical loads, must be cooled. Cooling of turbine blades is achieved by bleeding air from the compressor stage of the engine through complicated internal passages in the turbine blades (internal cooling, including jet-impingement cooling) and by bleeding small amounts of air into the boundary layer of the external flow through small discrete holes on the surface of the blade (film cooling and transpiration cooling). The cooling must be done using a minimum amount of air or any increases in efficiency gained through higher operating temperature will be lost due to added load on the compressor stage. Turbine cooling schemes have traditionally been based on extensive empirical data bases, quasi-one-dimensional computational fluid dynamics (CFD) analysis, and trial and error. With improved capabilities of CFD, these traditional methods can be augmented by full three-dimensional simulations of the coolant flow to predict in detail the heat transfer and metal temperatures. Several aspects of turbine coolant flows make such application of CFD difficult, thus a highly effective CFD methodology must be used. First, high resolution of the flow field is required to attain the needed accuracy for heat transfer predictions, making highly efficient flow solvers essential for such computations. Second, the geometries of the flow passages are complicated but must be modeled accurately in order to capture all important details of the flow. This makes grid generation and grid quality important issues. Finally, since coolant flows are turbulent and separated the effects of turbulence must be modeled with a low Reynolds number turbulence model to accurately predict details of heat transfer.

Rigby, David L.↗