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NASA NTRS2007

Testing of Flexible Ballutes in Hypersonic Wind Tunnels for Planetary Aerocapture

Studies were conducted for the In-Space Propulsion (ISP) Ultralightweight Ballute Technology Development Program to increase the technical readiness level of inflatable decelerator systems for planetary aerocapture. The present experimental study was conducted to develop the capability for testing lightweight, flexible materials in hypersonic facilities. The primary objectives were to evaluate advanced polymer film materials in a high-temperature, high-speed flow environment and provide experimental data for comparisons with fluid-structure interaction modeling tools. Experimental testing was conducted in the Langley Aerothermodynamics Laboratory 20-Inch Hypersonic CF4 and 31-Inch Mach 10 Air blowdown wind tunnels. Quantitative flexure measurements were made for 60 degree half angle afterbody-attached ballutes, in which polyimide films (1-mil and 3- mil thick) were clamped between a 1/2-inch diameter disk and a base ring (4-inch and 6-inch diameters). Deflection measurements were made using a parallel light silhouette of the film surface through an existing schlieren optical system. The purpose of this paper is to discuss these results as well as free-flying testing techniques being developed for both an afterbody-attached and trailing toroidal ballute configuration to determine dynamic fluid-structural stability. Methods for measuring polymer film temperature were also explored using both temperature sensitive paints (for up to 370 C) and laser-etched thin-film gages.

Buck, Gregory M.↗

NASA NTRS1987

A theoretical analysis of the extinction limits of a methane-air opposed-jet diffusion flame

A theoretical analysis is described for a methane-air diffusion flame stabilized in the forward stagnation region of a porous metal cylinder in a forced convective flow. The analysis includes effects of radiative heat loss from the porous metal surface and finite rate kinetics but neglects the effects of gravity. The theoretically predicted extinction limits compare well with experimentally observed extinction limits from the literature. After the predicted limits compared well with the experimental limits, a parametric study of the effect of fuel surface emissivity and Lewis number was conducted with the numerical model. It was found that the computed blowoff limit is independent of radiative heat loss for high fuel blowing velocities but is a strong function of Lewis number. At low fuel blowing velocities, the extinction limit varies with both radiative heat loss and Lewis number. It is discovered, however, that even if thermal losses from the fuel surface are absent, the flame can extinguish at the fuel surface independently of Lewis number due to excessive reaction zone thinning.

Olson, S. L.↗

NASA NTRS2009

NASA Tech Briefs, November 2009

Topics covered include: Cryogenic Chamber for Servo-Hydraulic Materials Testing; Apparatus Measures Thermal Conductance Through a Thin Sample from Cryogenic to Room Temperature; Rover Attitude and Pointing System Simulation Testbed; Desktop Application Program to Simulate Cargo-Air-Drop Tests; Multimodal Friction Ignition Tester; Small-Bolt Torque-Tension Tester; Integrated Spacesuit Audio System Enhances Speech Quality and Reduces Noise; Hardware Implementation of a Bilateral Subtraction Filter; Simple Optoelectronic Feedback in Microwave Oscillators; Small X-Band Oscillator Antennas; Free-Space Optical Interconnect Employing VCSEL Diodes; Discrete Fourier Transform Analysis in a Complex Vector Space; Miniature Scroll Pumps Fabricated by LIGA; Self-Assembling, Flexible, Pre-Ceramic Composite Preforms; Flight-speed Integral Image Analysis Toolkit; Work Coordination Engine; Multi-Mission Automated Task Invocation Subsystem; Autonomously Calibrating a Quadrupole Mass Spectrometer; Determining Spacecraft Reaction Wheel Friction Parameters; Composite Silica Aerogels Opacified with Titania; Multiplexed Colorimetric Solid-Phase Extraction; Detecting Airborne Mercury by Use of Polymer/Carbon Films; Lattice-Matched Semiconductor Layers on Single Crystalline Sapphire Substrate; Pressure-Energized Seal Rings to Better Withstand Flows; Rollerjaw Rock Crusher; Microwave Sterilization and Depyrogenation System; Quantifying Therapeutic and Diagnostic Efficacy in 2D Microvascular Images; NiF2/NaF:CaF2/Ca Solid-State High-Temperature Battery Cells; Critical Coupling Between Optical Fibers and WGM Resonators; Microwave Temperature Profiler Mounted in a Standard Airborne Research Canister; Alternative Determination of Density of the Titan Atmosphere; Solar Rejection Filter for Large Telescopes; Automated CFD for Generation of Airfoil Performance Tables; Progressive Classification Using Support Vector Machines; Active Learning with Irrelevant Examples; A Data Matrix Method for Improving the Quantification of Element Percentages of SEM/EDX Analysis; Deployable Shroud for the International X-Ray Observatory; Improved Model of a Mercury Ring Damper; Optoelectronic pH Meter: Further Details; X-38 Advanced Sublimator; and Solar Simulator Represents the Mars Surface Solar Environment.

Source record↗

DOE OSTI2020

Guided plasma jets directed onto wet surfaces: angular dependence and control

The optimal use of atmospheric pressure plasma jets (APPJs) for treatment of surfaces—inorganic, organic and liquid—depends on being able to control the flow of plasma-generated reactive species onto the surface. The typical APPJ is a rare gas mixture (RGM) flowed through a tube to which voltage is applied, producing an RGM plasma plume that extends into the ambient air. The RGM plasma plume is guided by a surrounding shroud of air due to the higher electric field required for an ionization wave (IW) to propagate into the air. Additionally, the mixing of the ambient air with the RGM plasma plume then determines the production of reactive oxygen and nitrogen species (RONS). The APPJ is usually oriented perpendicular to the surface being treated. However, the angle of the APPJ with respect to the surface may be a method to control the production of reactive species to the surface due to the change in APPJ propagation properties and the resulting gas dynamics. In this paper, we discuss results from computational and experimental investigations addressing two points—propagation of IWs in APPJs with and without a guiding gas shroud as a function of angle of the APPJ with respect to the surface; and the use of this angle to control plasma activation of thin water layers. We found that APPJs propagating out of the plasma tube into a same-gas environment lack any of the directional properties of shroud-guided jets, and largely follow electric field lines as the angle of the plasma tube is changed. Guided APPJs propagate coaxially with the tube as the angle is changed, and turn perpendicularly towards the surface only a few mm above the surface. The angle of the APPJ produces different gas dynamic distributions, which enable some degree of control over the content of RONS transferred to thin water layers.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS1996

Flow Coupling Effects in Jet-in-Crossflow Flowfields

The combustor designer is typically required to design liner orifices that effectively mix air jets with crossflow effluent. CFD combustor analysis is typically used in the design process; however the jets are usually assumed to enter the combustor with a uniform velocity and turbulence profile. The jet-mainstream flow coupling is usually neglected because of the computational expense. This CFD study was performed to understand the effect of jet-mainstream flow coupling, and to assess the accuracy of jet boundary conditions that are commonly used in combustor internal calculations. A case representative of a plenum-fed quick-mix section of a Rich Burn/Quick Mix/Lean Burn combustor (i.e. a jet-mainstream mass-flow ratio of about 3 and a jet-mainstream momentum-flux ratio of about 30) was investigated. This case showed that the jet velocity entering the combustor was very non-uniform, with a low normal velocity at the leading edge of the orifice and a high normal velocity at the trailing edge of the orifice. Three different combustor-only cases were analyzed with uniform inlet jet profile. None of the cases matched the plenum-fed calculations. To assess liner thickness effects, a thin-walled case was also analyzed. The CFD analysis showed the thin-walled jets had more penetration than the thick-walled jets.

Bain, D. B.↗

NASA NTRS2003

NASA Tech Briefs, September 2003

Topics include: Oxygen-Partial-Pressure Sensor for Aircraft Oxygen Mask; Three-Dimensional Venturi Sensor for Measuring Extreme Winds; Swarms of Micron-Sized Sensors; Monitoring Volcanoes by Use of Air-Dropped Sensor Packages; Capacitive Sensors for Measuring Masses of Cryogenic Fluids; UHF Microstrip Antenna Array for Synthetic- Aperture Radar; Multimode Broad-Band Patch Antennas; 164-GHz MMIC HEMT Frequency Doubler; GPS Position and Heading Circuitry for Ships; Software for Managing Parametric Studies; Software Aids Visualization of Computed Unsteady Flow; Software for Testing Electroactive Structural Components; Advanced Software for Analysis of High-Speed Rolling-Element Bearings; Web Program for Development of GUIs for Cluster Computers; XML-Based Generator of C++ Code for Integration With GUIs; Oxide Protective Coats for Ir/Re Rocket Combustion Chambers; Simplified Waterproofing of Aerogels; Improved Thermal-Insulation Systems for Low Temperatures; Device for Automated Cutting and Transfer of Plant Shoots; Extension of Liouville Formalism to Postinstability Dynamics; Advances in Thrust-Based Emergency Control of an Airplane; Ultrasonic/Sonic Mechanisms for Drilling and Coring; Exercise Device Would Exert Selectable Constant Resistance; Improved Apparatus for Measuring Distance Between Axles; Six Classes of Diffraction-Based Optoelectronic Instruments; Modernizing Fortran 77 Legacy Codes; Active State Model for Autonomous Systems; Shields for Enhanced Protection Against High-Speed Debris; Scaling of Two-Phase Flows to Partial-Earth Gravity; Neutral-Axis Springs for Thin-Wall Integral Boom Hinges.

Source record↗

NASA NTRS1986

Observations of field-aligned currents, waves, and electric fields at substorm onset

Substorm onsets, identified Pi 2 pulsations observed on the Air Force Geophysics Laboratory Magnetometer Network, are studied using magnetometer and electric field data from ISEE 1 as well as magnetometer data from the geosynchronous satellites GOES 2 and 3. The mid-latitude magnetometer data provides the means of both timing and locating the substorm onset so that the spacecraft locations with respect to the substorm current systems are known. During two intervals, each containing several onsets or intensifications, ISEE 1 observed field-aligned current signatures beginning simultaneously with the mid-latitude Pi 2 pulsation. Close to the earth broadband bursts of wave noise were observed in the electric field data whenever field-aligned currents were detected. One onset occurred when ISEE 1 and GOES 2 were on the same field line but in opposite hemispheres. During this onset ISEE 1 and GOES 2 saw magnetic signatures which appear to be due to conjugate field-aligned currents flowing out of the western end of the westward auroral electrojets. The ISEE 1 signature is of a line current moving westward past the spacecraft. During the other interval, ISEE 1 was in the near-tail region near the midnight meridian. Plasma data confirms that the plasma sheet thinned and subsequently expanded at onset. Electric field data shows that the plasma moved in the opposite direction to the plasma sheet boundary as the boundary expanded which implies that there must have been an abundant source of hot plasma present. The plasma motion was towards the center of the plasma sheet and earthwards and consisted of a series of pulses rather than a steady flow.

Smits, D. P.↗

NASA NTRS2024

The CapiSorb Visible System (CVS) Demonstrations on ISS

Falling liquid film amine sorbent reactors have been successfully employed to scrub CO 2 aboard submarines for decades. However, applying such proven methods aboard orbiting and coast spacecraft is significantly challenged by the nearly weightless environment, where liquid sprays and films do not fall, and vapor bubbles and gases do not rise. The Capillary Sorbent (CapiSorb) Visible System (CVS) is a technology demonstration experiment performed aboard the ISS April 18 – 21, 2023. The system establishes stable steady thin liquid film flows in Contactor (absorber) and Degasser (desorber/stripper) replacing the passive role of gravity with the combined passive roles of surface tension, wetting, and system geometry. A TOX-0 fructose ersatz liquid sorbent is employed enabling ‘transparent’ experiments performed and filmed by the crew safely in the open cabin of the ISS. Completed objectives include demonstrations of stable passive ‘massively’ parallel planar thin film capillary flows across atmospheric pressure Contactor and sealed heated Degasser. The impacts of varying flow rate, flow direction, heat input, viscosity, positive and negative Degasser pressures, condensate collection and return, fluid distribution, interfacial stability, and others are reported. At least 49 diagnostics are recorded for digitization and subsequent thermal-fluids model validation by a single HD video downlink during the nearly 22 hours of operations. An overview of the flight hardware including description of the components, diagnostics, crew procedures, flight operations, and summary of accomplishments is reported in Ref. 1. Further details of the diagnostics, tests performed, and data reduction is reported in Ref. 2. This report collects both1,2 into a single report adding methods of data digitization, reduction, and archive along with analyses and discussions of technology impacts.

microgravity↗

DOE OSTI2021

Performance Analysis and Limiting Parameters of Cross-flow Membrane-based Liquid-desiccant Air Dehumidifiers

We report that to dehumidify a humid air stream, existing air conditioning (AC) systems substantially overcool the outdoor humid air below its dew point, thereby significantly reducing energy efficiency. Directly capturing humidity, membrane-based liquid-desiccant dehumidification systems separate sensible and latent cooling (SSLC) loads and thus offer a promising pathway for a high-performance AC solution. Design of an energy-efficient SSLC-AC system, however, rests largely on detailed understating of the dehumidification process. While some studies have identified the dehumidification process mainly depends on membrane characteristics, other studies have argued that the process is limited by desiccant liquid or alternatively air thermo-hydraulic physics for typical humid climate conditions. The present study examines performance and physics of the membrane-based liquid-desiccant dehumidification process over a wide range of climate conditions through a novel 3D, two-phase, multi-species CFD model. Decoupling the thermodynamic and hydraulic effects, the study reveals that the dehumidification rate is a linear function of the water vapor pressure potential ($J=α ΔP$) summarizing the system's thermodynamic state. The slope of the curve (i.e., α) depends on hydraulic transport characteristics of the membrane pores, air stream, and desiccant solution. More importantly, it was found that the air dehumidification process is mainly limited by the air-side transport physics for thin liquid-desiccant films and commonly used porous superhydrophobic membranes. Additionally, results show that, depending on ambient/desiccant conditions and physical dehumidifier characteristics, energy effectiveness and dehumidification rate vary from 13 to 34% and from 0.13 to 1.4 g m -2 s -1 , respectively. Therefore, the present study allows to efficiently design future SSLC-based AC systems exhibiting high performance energy metrics.

29 ENERGY PLANNING, POLICY, AND ECONOMY↗

NASA NTRS2016

Design and Fabrication of Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged Combustion Thrust Chamber Injectors

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. On the current project, several configurations of new main injectors were considered for the thrust chamber assembly of the integrated test article. All the injector elements were of the gas-centered swirl coaxial type, similar to those used on the Russian oxidizer-rich staged-combustion rocket engines. In such elements, oxidizer-rich combustion products from the preburner/turbine exhaust flow through a straight tube, and fuel exiting from the combustion chamber and nozzle regenerative cooling circuits is injected near the exit of the oxidizer tube through tangentially oriented orifices that impart a swirl motion such that the fuel flows along the wall of the oxidizer tube in a thin film. In some elements there is an orifice at the inlet to the oxidizer tube, and in some elements there is a sleeve or "shield" inside the oxidizer tube where the fuel enters. In the current project, several variations of element geometries were created, including element size (i.e., number of elements or pattern density), the distance from the exit of the sleeve to the injector face, the width of the gap between the oxidizer tube inner wall and the outer wall of the sleeve, and excluding the sleeve entirely. This paper discusses the design rationale for each of these element variations, including hydraulic, structural, thermal, combustion performance, and combustion stability considerations. This paper also discusses the fabrication and assembly of the injector components, including the injector body/interpropellant plate, the additive manufactured GRCop-84 faceplate, and the pieces that make up the injector elements including the oxidizer tube, an inlet to the oxidizer tube, and a facenut that includes the fuel tangential inlets and forms the initial recessed volume where oxidizer and fuel first interact. Hot-fire test results of these main injector designs in an integrated test article that includes an oxidizer-rich preburner are described in companion papers at this JANNAF meeting.

Garcia, C. P.↗

NASA NTRS1997

Experimental and Computational Studies of the Flow Over a Sting Mounted Planetary Probe Configuration

This paper summarizes the results of a series of experimental studies in the LENS shock tunnel and computations with DSMC and Navier Stokes codes which have been made to examine the aerothermal and flowfield characteristics of the flow over a sting-supported planetary probe configuration in hypervelocity air and nitrogen flows. The experimental program was conducted in the LENS hypervelocity shock tunnel at total enthalpies of 5and 10 MJkg for a range of reservoir pressure conditions from 70 to 500 bars. Heat transfer and pressure measurements were made on the front and rear face of the probe and along the supporting sting. High-speed and single shot schlieren photography were also employed to examine the flow over the model and the time to establish the flow in the base recirculation region. Predictions of the flowfield characteristics and the distributions of heat transfer and pressure were made with DSMC codes for rarefied flow conditions and with the Navier-Stokes solvers for the higher pressure conditions where the flows were assumed to be laminar. Analysis of the time history records from the heat transfer and pressure instrumentation on the face of the probe and in the base region indicated that the base flow was fully established in under 4 milliseconds from flow initiation or between 35 and 50 flow lengths based on base height. The measurements made in three different tunnel entries with two models of identical geometries but with different instrumentation packages, one prepared by NASA Langley and the second prepared by CUBRC, demonstrated good agreement between heat transfer measurements made with two different types of thin film and coaxial gage instrumentation. The measurements of heat transfer and pressure to the front face of the probe were in good agreement with theoretical predictions from both the DSMC and Navier Stokes codes. For the measurements made in low density flows, computations with the DSMC code were found to compare well with the pressure and heat transfer measurements on the sting, although the computed heat transfer rates in the recirculation region did not exhibit the same characteristics as the measurements. For the 10MJkg and 500 bar reservoir match point condition, the measurements and heat transfer along the sting from the first group of studies were in agreement with the Navier Stokes solutions for laminar conditions. A similar set of measurements made in later tests where the model was moved to a slightly different position in the test section indicated that the boundary layer in the reattachment compression region was close to transition or transitional where small changes in the test environment can result in larger than laminar heating rates. The maximum heating coefficients on the sting observed in the present studies was a small fraction of similar measurements obtained at nominally the same conditions in the HEG shock tunnel, where it is possible for transition to occur in the base flow, and in the low enthalpy studies conducted in the NASA Langley high Reynolds number Mach 10 tunnel where the base flow was shown to be turbulent. While the hybrid Navier- StokedDMSC calculations by Gochberg et al. (Reference 1) suggested that employing the Navier- Stokes calculations for the entire flowfield could be seriously in error in the base region for the 10 MJkg, 500 bar test case, similar calculations performed by Cornell, presented here, do not.

Holden, Michael S.↗

NASA NTRS1995

Thrust measurement in a 2-D scramjet nozzle

The two-dimensional thrust nozzle presents a challenging problem. The loading is not axisymmetric as in the case of a cone and the internal flow presents some design difficulties. A two-sting system has been chosen to accomodate the internal flow and achieve some symmetry. The situation is complicated by the fact that with the small ramp angle and the internal pressure on the nozzle walls, loading is predominantly transverse. Yet it is the axial thrust which is to be measured (i.e., the tensile waves propagating in the stings). Although bending stress waves travel at most at only 60% of the speed of the axial stress waves, the system needs to be stiffened against bending. The second sting was originally only used to preserve symmetry. However, the pressures on each thrust surface may be quite different at some conditions, so at this stage the signals from both stings are being averaged as a first order approximation of the net thrust. The expected axial thrust from this nozzle is not large so thin stings are required. In addition, the contact area between nozzle and sting needs to be maximized. The result was that it was decided to twist the stings through 90 deg, without distorting their cross-sectional shape, just aft of the nozzle. Finite element analysis showed that this would not significantly alter the propagation of the axial stress wave in the sting, while the rigidity of the system is greatly increased. A Mach 4 contoured nozzle is used in the experiments. The thrust calculated by integrating the static pressure measurements on the thrust surfaces is compared with the deconvolved strain measurement of the net thrust for the cases of air only and hydrogen fuel injected into air at approximately 9 MJ/kg nozzle supply enthalpy. The gain in thrust due to combustion is visible in this result.

Tuttle, Sean↗

NASA NTRS1999

NASA GRC's High Pressure Burner Rig Facility and Materials Test Capabilities

The High Pressure Burner Rig (HPBR) at NASA Glenn Research Center is a high-velocity. pressurized combustion test rig used for high-temperature environmental durability studies of advanced materials and components. The facility burns jet fuel and air in controlled ratios, simulating combustion gas chemistries and temperatures that are realistic to those in gas turbine engines. In addition, the test section is capable of simulating the pressures and gas velocities representative of today's aircraft. The HPBR provides a relatively inexpensive. yet sophisticated means for researchers to study the high-temperature oxidation of advanced materials. The facility has the unique capability of operating under both fuel-lean and fuel-rich gas mixtures. using a fume incinerator to eliminate any harmful byproduct emissions (CO, H2S) of rich-burn operation. Test samples are easily accessible for ongoing inspection and documentation of weight change, thickness, cracking, and other metrics. Temperature measurement is available in the form of both thermocouples and optical pyrometery. and the facility is equipped with quartz windows for observation and video taping. Operating conditions include: (1) 1.0 kg/sec (2.0 lbm/sec) combustion and secondary cooling airflow capability: (2) Equivalence ratios of 0.5- 1.0 (lean) to 1.5-2.0 (rich), with typically 10% H2O vapor pressure: (3) Gas temperatures ranging 700-1650 C (1300-3000 F): (4) Test pressures ranging 4-12 atmospheres: (5) Gas flow velocities ranging 10-30 m/s (50-100) ft/sec.: and (6) Cyclic and steady-state exposure capabilities. The facility has historically been used to test coupon-size materials. including metals and ceramics. However complex-shaped components have also been tested including cylinders, airfoils, and film-cooled end walls. The facility has also been used to develop thin-film temperature measurement sensors.

Robinson, R. Craig↗

NASA NTRS1999

High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure: Experimental and Analytical Validation

Experimental data were obtained on an optimally contoured nozzle with an area ratio of 1025:1 and on a truncated version of this nozzle with an area ratio of 440:1. The nozzles were tested with gaseous hydrogen and liquid oxygen propellants at combustion chamber pressures of 1800 to 2400 psia and mixture ratios of 3.89 to 6.15. This report compares the experimental performance, heat transfer, and boundary layer total pressure measurements with theoretical predictions of the current Joint Army, Navy, NASA, Air Force (JANNAF) developed methodology. This methodology makes use of the Two-Dimensional Kinetics (TDK) nozzle performance code. Comparisons of the TDK-predicted performance to experimentally attained thrust performance indicated that both the vacuum thrust coefficient and the vacuum specific impulse values were approximately 2.0-percent higher than the turbulent prediction for the 1025:1 configurations, and approximately 0.25-percent higher than the turbulent prediction for the 440:1 configuration. Nozzle wall temperatures were measured on the outside of a thin-walled heat sink nozzle during the test fittings. Nozzle heat fluxes were calculated front the time histories of these temperatures and compared with predictions made with the TDK code. The heat flux values were overpredicted for all cases. The results range from nearly 100 percent at an area ratio of 50 to only approximately 3 percent at an area ratio of 975. Values of the integral of the heat flux as a function of nozzle surface area were also calculated. Comparisons of the experiment with analyses of the heat flux and the heat rate per axial length also show that the experimental values were lower than the predicted value. Three boundary layer rakes mounted on the nozzle exit were used for boundary layer measurements. This arrangement allowed total pressure measurements to be obtained at 14 different distances from the nozzle wall. A comparison of boundary layer total pressure profiles and analytical predictions show good agreement for the first 0.5 in. from the nozzle wall; but the further into the core flow that measurements were taken, the more that TDK overpredicted the boundary layer thickness.

Jankovsky, Robert S.↗

DOE OSTI2020

Steady detonation propagation in thin channels with strong confinement

We examine asymptotically the dynamics of two-dimensional, steady detonation wave propagation and failure for a strongly confined high explosive (HE), in which the width of the explosive is small relative to the reaction zone length. An energy balance equation is derived, which shows how the longitudinal acceleration of subsonic flow behind the detonation shock is influenced both by chemical reaction and by the effects of HE boundary streamline deflection, specifically via the induced rate of change of mass flux through the detonation wave. The latter serves to either counteract or reinforce the acceleration of longitudinal flow, depending on the sign of the gradient of the boundary streamline deflection at the detonation shock. The analysis is valid for general equations of state and chemical reaction rates in the HE. The asymptotically derived form of the energy equation represents an eigenvalue problem for the determination of the steady detonation propagation speed, solved via a shooting method. We explore specific results for ideal and stiffened equations of state, along with a pressure-dependent reaction rate for which changes in the pressure exponent and reaction order are also studied. We consider the influences of both straight and curved HE boundary streamline shapes. The asymptotic analysis reveals significant physical insights into how detonation propagation and failure are affected by strong confinement.

42 ENGINEERING↗

NASA NTRS1998

Bulk-Flow Analysis of Hybrid Thrust Bearings for Advanced Cryogenic Turbopumps

A bulk-flow analysis and computer program for prediction of the static load performance and dynamic force coefficients of angled injection, orifice-compensated hydrostatic/hydrodynamic thrust bearings have been completed. The product of the research is an efficient computational tool for the design of high-speed thrust bearings for cryogenic fluid turbopumps. The study addresses the needs of a growing technology that requires of reliable fluid film bearings to provide the maximum operating life with optimum controllable rotordynamic characteristics at the lowest cost. The motion of a cryogenic fluid on the thin film lands of a thrust bearing is governed by a set of bulk-flow mass and momentum conservation and energy transport equations. Mass flow conservation and a simple model for momentum transport within the hydrostatic bearing recesses are also accounted for. The bulk-flow model includes flow turbulence with fluid inertia advection, Coriolis and centrifugal acceleration effects on the bearing recesses and film lands. The cryogenic fluid properties are obtained from realistic thermophysical equations of state. Turbulent bulk-flow shear parameters are based on Hirs' model with Moody's friction factor equations allowing a simple simulation for machined bearing surface roughness. A perturbation analysis leads to zeroth-order nonlinear equations governing the fluid flow for the thrust bearing operating at a static equilibrium position, and first-order linear equations describing the perturbed fluid flow for small amplitude shaft motions in the axial direction. Numerical solution to the zeroth-order flow field equations renders the bearing flow rate, thrust load, drag torque and power dissipation. Solution to the first-order equations determines the axial stiffness, damping and inertia force coefficients. The computational method uses well established algorithms and generic subprograms available from prior developments. The Fortran9O computer program hydrothrust runs on a Windows 95/NT personal computer. The program, help files and examples are licensed by Texas A&M University Technology License Office. The study of the static and dynamic performance of two hydrostatic/hydrodynamic bearings demonstrates the importance of centrifugal and advection fluid inertia effects for operation at high rotational speeds. The first example considers a conceptual hydrostatic thrust bearing for an advanced liquid hydrogen turbopump operating at 170,000 rpm. The large axial stiffness and damping coefficients of the bearing should provide accurate control and axial positioning of the turbopump and also allow for unshrouded impellers, therefore increasing the overall pump efficiency. The second bearing uses a refrigerant R134a, and its application in oil-free air conditioning compressors is of great technological importance and commercial value. The computed predictions reveal that the LH2 bearing load capacity and flow rate increase with the recess pressure (i.e. increasing orifice diameters). The bearing axial stiffness has a maximum for a recess pressure rati of approx. 0.55. while the axial damping coefficient decreases as the recess pressure ratio increases. The computer results from three flow models are compared. These models are a) inertialess, b) fluid inertia at recess edges only, and c) full fluid inertia at both recess edges and film lands. The full inertia model shows the lowest flow rates, axial load capacity and stiffness coefficient but on the other hand renders the largest damping coefficients and inertia coefficients. The most important findings are related to the reduction of the outflow through the inner radius and the appearance of subambient pressures. The performance of the refrigerant hybrid thrust bearing is evaluated at two operating speeds and pressure drops. The computed results are presented in dimensionless form to evidence consistent trends in the bearing performance characteristics. As the applied axial load increases, the bearing film thickness and flow rate decrease while the recess pressure increases. The axial stiffness coefficient shows a maximum for a certain intermediate load while the damping coefficient steadily increases. The computed results evidence the paramount of centrifugal fluid inertia at low recess pressures (i.e. low loads), and where there is actually an inflow through the bearing inner diameter, accompanied by subambient pressures just downstream of the bearing recess edge. These results are solely due to centrifugal fluid inertia and advection transport effects. Recommendations include the extension of the computer program to handle flexure pivot tilting pad hybrid bearings and the ability to calculate moment coefficients for shaft angular misalignments.

SanAndres, Luis↗

NASA NTRSDate not supplied

Integral Channel Nozzles and Heat Exchangers using Additive Manufacturing Directed Energy Deposition NASA HR-1 Alloy

Heat exchangers for use in propulsion applications are very critical components because they must be efficient, compact and light and often operate with working fluids at extreme temperatures or pressures or both. Various components and systems use heat exchangers such as combustion chambers of gas turbines and internal combustion engines, fuel cells (air supply and thermal management), electric batteries (thermal management), evaporators and recuperators of waste-heat-to-power systems, and rocket engines. Even if the results are more generally applicable, the heat exchangers applications to which this study is more closely related are regeneratively cooled rocket nozzles and chambers, and repressurization systems for the launch vehicles. These components are often thin-walled and contain pressurized fluids, like propellants at cryogenic or elevated temperatures. Given that the environments that these propulsion components must endure are challenging, the manufacturing to meet these specifications often require long lead times due to specialty processes and unique tooling associated with the combined thin-wall integral channel and large-scale structures. Additive manufacturing (AM) offers programmatic advantages for reduction in processing time and cost in addition to various technical advantages, including the possibility to achieve enhanced hardware complexity targeted to superior performance, part consolidation, and the capability of processing of novel alloys. While AM is already being utilized for heat exchanger components in propulsion applications, almost all these AM components are made by means of Laser Powder Bed Fusion (L-PBF). L-PBF allows for fine features but is rather limited with respect to the overall size of the components that can be manufactured. Recent developments are maturing the Laser Powder Directed Energy Deposition (LP-DED) process which may be used, for example, to make integral channel thin-wall regeneratively-cooled rocket nozzles with diameters greater than 1 m. This paper highlights some integral channel heat exchanger demonstrator hardware applications of LP-DED, as well as the characterization of this process in combination with the use of the NASA HR-1 alloy. To properly utilize LP-DED for heat exchanger manufacturing, various aspects are being characterized such as geometry limitations, measurement of surface texture and geometric angled surfaces, surface enhancements for internal channels, and material evaluation. NASA HR-1 (FeNi-Cr) is a high strength hydrogen resistant superalloy developed for use in aerospace applications, such as heat exchangers. Some aspects and considerations about the design of heat exchangers are summarized together with data relevant to LP-DED manufacturing in combination with the NASA HR-1 alloy. Microchannels were successful deposited down to 2.54 mm and 1 mm wall thickness, wall angles of 30°, both with high reproducibility. It was also found that the areal surface roughness is highly dependent on the size of the powder feedstock used for deposition. The characterization of these LP-DED features is critical for fluid flow and heat transfer predictions as it can be exploited to enhance heat transfer at the cost of increased pressure drop.

Additive Manufacturing↗

NASA NTRSDate not supplied

Integral Channel Nozzles and Heat Exchangers using Additive Manufacturing Directed Energy Deposition NASA HR-1 Alloy

additive manufacturing↗