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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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NASA NTRS2007

Discrete Roughness Transition for Hypersonic Flight Vehicles

The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.

Berry, Scott A.↗
NASA NTRS2007

Phoenix Missile Hypersonic Testbed (PMHT): Project Concept Overview

An over view of research into a low cost hypersonic research flight test capability to increase the amount of hypersonic flight data to help bridge the large developmental gap between ground testing/analysis and major flight demonstrator Xplanes is provided. The major objectives included: develop an air launched missile booster research testbed; accurately deliver research payloads through programmable guidance to hypersonic test conditions; low cost; a high flight rate minimum of two flights per year and utilize surplus air launched missiles and NASA aircraft.

Jones, Thomas P.↗
NASA NTRS2007

Simulation of Stagnation Region Heating in Hypersonic Flow on Tetrahedral Grids

Hypersonic flow simulations using the node based, unstructured grid code FUN3D are presented. Applications include simple (cylinder) and complex (towed ballute) configurations. Emphasis throughout is on computation of stagnation region heating in hypersonic flow on tetrahedral grids. Hypersonic flow over a cylinder provides a simple test problem for exposing any flaws in a simulation algorithm with regard to its ability to compute accurate heating on such grids. Such flaws predominantly derive from the quality of the captured shock. The importance of pure tetrahedral formulations are discussed. Algorithm adjustments for the baseline Roe / Symmetric, Total-Variation-Diminishing (STVD) formulation to deal with simulation accuracy are presented. Formulations of surface normal gradients to compute heating and diffusion to the surface as needed for a radiative equilibrium wall boundary condition and finite catalytic wall boundary in the node-based unstructured environment are developed. A satisfactory resolution of the heating problem on tetrahedral grids is not realized here; however, a definition of a test problem, and discussion of observed algorithm behaviors to date are presented in order to promote further research on this important problem.

Gnoffo, Peter A.↗
NASA NTRS2008

Analysis of the Effects of Vitiates on Surface Heat Flux in Ground Tests of Hypersonic Vehicles

To achieve the high enthalpy conditions associated with hypersonic flight, many ground test facilities burn fuel in the air upstream of the test chamber. Unfortunately, the products of combustion contaminate the test gas and alter gas properties and the heat fluxes associated with aerodynamic heating. The difference in the heating rates between clean air and a vitiated test medium needs to be understood so that the thermal management system for hypersonic vehicles can be properly designed. This is particularly important for advanced hypersonic vehicle concepts powered by air-breathing propulsion systems that couple cooling requirements, fuel flow rates, and combustor performance by flowing fuel through sub-surface cooling passages to cool engine components and preheat the fuel prior to combustion. An analytical investigation was performed comparing clean air to a gas vitiated with methane/oxygen combustion products to determine if variations in gas properties contributed to changes in predicted heat flux. This investigation started with simple relationships, evolved into writing an engineering-level code, and ended with running a series of CFD cases. It was noted that it is not possible to simultaneously match all of the gas properties between clean and vitiated test gases. A study was then conducted selecting various combinations of freestream properties for a vitiated test gas that matched clean air values to determine which combination of parameters affected the computed heat transfer the least. The best combination of properties to match was the free-stream total sensible enthalpy, dynamic pressure, and either the velocity or Mach number. This combination yielded only a 2% difference in heating. Other combinations showed departures of up to 10% in the heat flux estimate.

Cuda, Vincent↗
NASA NTRS2007

Discrete Roughness Transition for Hypersonic Flight Vehicles

The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.

Berry, Scott A.↗
NASA NTRS2008

Stereoscopic Imaging in Hypersonics Boundary Layers using Planar Laser-Induced Fluorescence

Stereoscopic time-resolved visualization of three-dimensional structures in a hypersonic flow has been performed for the first time. Nitric Oxide (NO) was seeded into hypersonic boundary layer flows that were designed to transition from laminar to turbulent. A thick laser sheet illuminated and excited the NO, causing spatially-varying fluorescence. Two cameras in a stereoscopic configuration were used to image the fluorescence. The images were processed in a computer visualization environment to provide stereoscopic image pairs. Two methods were used to display these image pairs: a cross-eyed viewing method which can be viewed by naked eyes, and red/blue anaglyphs, which require viewing through red/blue glasses. The images visualized three-dimensional information that would be lost if conventional planar laser-induced fluorescence imaging had been used. Two model configurations were studied in NASA Langley Research Center's 31-Inch Mach 10 Air Wind tunnel. One model was a 10 degree half-angle wedge containing a small protuberance to force the flow to transition. The other model was a 1/3-scale, truncated Hyper-X forebody model with blowing through a series of holes to force the boundary layer flow to transition to turbulence. In the former case, low flowrates of pure NO seeded and marked the boundary layer fluid. In the latter, a trace concentration of NO was seeded into the injected N2 gas. The three-dimensional visualizations have an effective time resolution of about 500 ns, which is fast enough to freeze this hypersonic flow. The 512x512 resolution of the resulting images is much higher than high-speed laser-sheet scanning systems with similar time response, which typically measure 10-20 planes.

Danehy, Paul M.↗
NASA NTRS2009

Compressibility Considerations for kappa-omega Turbulence Models in Hypersonic Boundary Layer Applications

The ability of kappa-omega models to predict compressible turbulent skin friction in hypersonic boundary layers is investigated. Although uncorrected two-equation models can agree well with correlations for hot-wall cases, they tend to perform progressively worse - particularly for cold walls - as the Mach number is increased in the hypersonic regime. Simple algebraic models such as Baldwin-Lomax perform better compared to experiments and correlations in these circumstances. Many of the compressibility corrections described in the literature are summarized here. These include corrections that have only a small influence for kappa-omega models, or that apply only in specific circumstances. The most widely-used general corrections were designed for use with jet or mixing-layer free shear flows. A less well-known dilatation-dissipation correction intended for boundary layer flows is also tested, and is shown to agree reasonably well with the Baldwin-Lomax model at cold-wall conditions. It exhibits a less dramatic influence than the free shear type of correction. There is clearly a need for improved understanding and better overall physical modeling for turbulence models applied to hypersonic boundary layer flows.

Rumsey, C. L.↗
NASA NTRS2009

Hypersonic Navier Stokes Comparisons to Orbiter Flight Data

Hypersonic chemical nonequilibrium simulations of low earth orbit entry flow fields are becoming increasingly commonplace as software and computational capabilities become more capable. However, development of robust and accurate software to model these environments will always encounter a significant barrier in developing a suite of high quality calibration cases. The US3D hypersonic nonequilibrium Navier Stokes analysis capability has been favorably compared to a number of wind tunnel test cases. Extension of the calibration basis for this software to Orbiter flight conditions will provide an incremental increase in confidence. As part of the Orbiter Boundary Layer Transition Flight Experiment and the Hypersonic Thermodynamic Infrared Measurements project, NASA is performing entry flight testing on the Orbiter to provide valuable aerothermodynamic heating data. An increase in interest related to orbiter entry environments is resulting from this activity. With the advent of this new data, comparisons of the US3D software to the new flight testing data is warranted. This paper will provide information regarding the framework of analyses that will be applied with the US3D analysis tool. In addition, comparisons will be made to entry flight testing data provided by the Orbiter BLT Flight Experiment and HYTHIRM projects. If data from digital scans of the Orbiter windward surface become available, simulations will also be performed to characterize the difference in surface heating between the CAD reference OML and the digitized surface provided by the surface scans.

Campbell, Charles H.↗
NASA NTRS2009

An Overview of the NASA FAP Hypersonics Project Airbreathing Propulsion Research

The propulsion research portfolio of the National Aeronautics and Space Administration Fundamental Aeronautics Program Hypersonics Project encompasses a significant number of technical tasks that are aligned to achieve mastery and intellectual stewardship of the core competencies in the hypersonic-flight regime. An overall coordinated programmatic and technical effort has been structured to advance the state-of-the-art, via both experimental and analytical efforts. A subset of the entire hypersonics propulsion research portfolio is presented in this overview paper. To this end, two programmatic research disciplines are discussed; namely, (1) the Propulsion Discipline, including three associated research elements: the X-51A partnership, the HIFiRE-2 partnership, and the Durable Combustor Rig, and (2) the Turbine-Based Combine Cycle Discipline, including three associated research elements: the Combined Cycle Engine Large Scale Inlet Mode Transition Experiment, the small-scale Inlet Mode Transition Experiment, and the High-Mach Fan Rig.

Auslender, A. H.↗
NASA NTRS2010

The HYTHIRM Project: Flight Thermography of the Space Shuttle During the Hypersonic Re-entry

This report describes a NASA Langley led endeavor sponsored by the NASA Engineering Safety Center, the Space Shuttle Program Office and the NASA Aeronautics Research Mission Directorate to demonstrate a quantitative thermal imaging capability. A background and an overview of several multidisciplinary efforts that culminated in the acquisition of high resolution calibrated infrared imagery of the Space Shuttle during hypervelocity atmospheric entry is presented. The successful collection of thermal data has demonstrated the feasibility of obtaining remote high-resolution infrared imagery during hypersonic flight for the accurate measurement of surface temperature. To maximize science and engineering return, the acquisition of quantitative thermal imagery and capability demonstration was targeted towards three recent Shuttle flights - two of which involved flight experiments flown on Discovery. In coordination with these two Shuttle flight experiments, a US Navy NP-3D aircraft was flown between 26-41 nautical miles below Discovery and remotely monitored surface temperature of the Orbiter at Mach 8.4 (STS-119) and Mach 14.7 (STS-128) using a long-range infrared optical package referred to as Cast Glance. This same Navy aircraft successfully monitored the Orbiter Atlantis traveling at approximately Mach 14.3 during its return from the successful Hubble repair mission (STS-125). The purpose of this paper is to describe the systematic approach used by the Hypersonic Thermodynamic Infrared Measurements team to develop and implement a set of mission planning tools designed to establish confidence in the ability of an imaging platform to reliably acquire, track and return global quantitative surface temperatures of the Shuttle during entry. The mission planning tools included a pre-flight capability to predict the infrared signature of the Shuttle. Such tools permitted optimization of the hardware configuration to increase signal-to-noise and to maximize the available dynamic range while mitigating the potential for saturation. Post flight, analysis tools were used to assess atmospheric effects and to convert the 2-D intensity images to 3-D temperature maps of the windward surface. Comparison of the spatially resolved global thermal measurements to surface thermocouples and CFD prediction is made. Successful demonstration of a quantitative, spatially resolved, global temperature measurement on the Shuttle suggests future applications towards hypersonic flight test programs within NASA, DoD and DARPA along with flight test opportunities supporting NASA's project Constellation.

Horvath, Thomas J.↗
NASA NTRS2011

Dynamic Testing of the NASA Hypersonic Project Combined Cycle Engine Testbed for Mode Transition Experiments

NASA is interested in developing technology that leads to more routine, safe, and affordable access to space. Access to space using airbreathing propulsion systems has potential to meet these objectives based on Airbreathing Access to Space (AAS) system studies. To this end, the NASA Fundamental Aeronautics Program (FAP) Hypersonic Project is conducting fundamental research on a Turbine Based Combined Cycle (TBCC) propulsion system. The TBCC being studied considers a dual flow-path inlet system. One flow-path includes variable geometry to regulate airflow to a turbine engine cycle. The turbine cycle provides propulsion from take-off to supersonic flight. The second flow-path supports a dual-mode scramjet (DMSJ) cycle which would be initiated at supersonic speed to further accelerate the vehicle to hypersonic speed. For a TBCC propulsion system to accelerate a vehicle from supersonic to hypersonic speed, a critical enabling technology is the ability to safely and effectively transition from the turbine to the DMSJ-referred to as mode transition. To experimentally test methods of mode transition, a Combined Cycle Engine (CCE) Large-scale Inlet testbed was designed with two flow paths-a low speed flow-path sized for a turbine cycle and a high speed flow-path designed for a DMSJ. This testbed system is identified as the CCE Large-Scale Inlet for Mode Transition studies (CCE-LIMX). The test plan for the CCE-LIMX in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel (10x10 SWT) is segmented into multiple phases. The first phase is a matrix of inlet characterization (IC) tests to evaluate the inlet performance and establish the mode transition schedule. The second phase is a matrix of dynamic system identification (SysID) experiments designed to support closed-loop control development at mode transition schedule operating points for the CCE-LIMX. The third phase includes a direct demonstration of controlled mode transition using a closed loop control system developed with the data obtained from the first two phases. Plans for a fourth phase include mode transition experiments with a turbine engine. This paper, focusing on the first two phases of experiments, presents developed operational and analysis tools for streamlined testing and data reduction procedures.

Source record↗
NASA NTRS2012

Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths

This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.

Georgiadis, Nicholas J.↗
NASA NTRS1959

Experimental Verification of Boundary-Layer Corrections in Hypersonic Nozzles

The problem of accurately predicting boundary-layer growth for supersonic nozzle design assumes increased importance as the design Mach Number of the nozzles is pushed into the hypersonic regime. Several methods of calculating this growth have been advanced which vary in ease of computation and accuracy of results. Results of the calibrations performed on two hypersonic helium nozzles recently put into operation in the 11-in. Hypersonic Tunnel Section at the NASA Langley Research Center indicate that the method employed for boundary-layer calculations in these nozzles is adequate. The flow in the test region of both nozzles has been found to be of good quality and the Mach Numbers obtained agree very satisfactorily with the desired design Mach Number.

Boundary Layer Equations↗
NASA NTRS2015

An Upgrade of the Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Software

The Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) code is used at NASA Langley Research Center to analyze global aeroheating data on wind tunnel models tested in the Langley Aerothermodynamics Laboratory. One-dimensional, semi-infinite heating data derived from IHEAT are used to design thermal protection systems to mitigate the risks due to the aeroheating loads on hypersonic vehicles, such as re-entry vehicles during descent and landing procedures. This code was originally written in the PV-WAVE programming language to analyze phosphor thermography data from the two-color, relativeintensity system developed at Langley. To increase the efficiency, functionality, and reliability of IHEAT, the code was migrated to MATLAB syntax and compiled as a stand-alone executable file labeled version 4.0. New features of IHEAT 4.0 include the options to batch process all of the data from a wind tunnel run, to map the two-dimensional heating distribution to a three-dimensional computer-aided design model of the vehicle to be viewed in Tecplot, and to extract data from a segmented line that follows an interesting feature in the data. Results from IHEAT 4.0 were compared on a pixel level to the output images from the legacy code to validate the program. The differences between the two codes were on the order of 10-5 to 10-7. IHEAT 4.0 replaces the PV-WAVE version as the production code for aeroheating experiments conducted in the hypersonic facilities at NASA Langley.

Mason, Michelle L.↗
NASA NTRS2016

Features of the Upgraded Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Software

The Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) software is used at the NASA Langley Research Center to analyze global aeroheating data on wind tunnel models tested in the Langley Aerothermodynamics Laboratory. One-dimensional, semi-infinite heating data derived from IHEAT are used in the design of thermal protection systems for hypersonic vehicles that are exposed to severe aeroheating loads, such as reentry vehicles during descent and landing procedures. This software program originally was written in the PV-WAVE(Registered Trademark) programming language to analyze phosphor thermography data from the two-color, relative-intensity system developed at Langley. To increase the efficiency, functionality, and reliability of IHEAT, the program was migrated to MATLAB(Registered Trademark) syntax and compiled as a stand-alone executable file labeled version 4.0. New features of IHEAT 4.0 include the options to perform diagnostic checks of the accuracy of the acquired data during a wind tunnel test, to extract data along a specified multi-segment line following a feature such as a leading edge or a streamline, and to batch process all of the temporal frame data from a wind tunnel run. Results from IHEAT 4.0 were compared on a pixel level to the output images from the legacy software to validate the program. The absolute differences between the heat transfer data output from the two programs were on the order of 10(exp -5) to 10(exp -7). IHEAT 4.0 replaces the PV-WAVE(Registered Trademark) version as the production software for aeroheating experiments conducted in the hypersonic facilities at NASA Langley.

Mason, Michelle L.↗
NASA NTRS2018

Characterization of Freestream Disturbances in Conventional Hypersonic Wind Tunnels

While low disturbance ("quiet") hypersonic wind tunnels are believed to provide more reliable extrapolation of boundary layer transition behavior from ground to flight, the presently available quiet facilities are limited to Mach 6, moderate Reynolds numbers, low freestream enthalpy, and subscale models. As a result, only conventional ("noisy") wind tunnels can reproduce both Reynolds numbers and enthalpies of hypersonic flight configurations, and must therefore be used for flight vehicle test and evaluation involving high Mach number, high enthalpy, and larger models. This article outlines the recent progress and achievements in the characterization of tunnel noise that have resulted from the coordinated effort within the AVT-240 specialists group on hypersonic boundary layer transition prediction. New Direct Numerical Simulation (DNS) datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the spatiotemporal structure of the freestream noise, and account for the propagation and transfer of the freestream disturbances to a pitot-mounted sensor. The new experimental measurements cover a range of conventional wind tunnels with different sizes and Mach numbers from 6 to 14 and extend the database of freestream fluctuations within the spectral range of boundary layer instability waves over commonly tested models. Prospects for applying the computational and measurement datasets for developing mechanism-based transition prediction models are discussed.

Duan, Lian↗
NASA NTRS2018

Hypersonic Chemically Reacting Boundary-Layer Stability using LASTRAC

Hypersonic boundary layer transition is critical to the design of all hypersonic vehicles due to its effect on the heat transfer into the vehicle surface and potential drag enhancement or reduction during reentry. Boundary layer transition and boundary layer stability analysis under hypersonic conditions has been studied for decades, yet there is ample room for improved accuracy and further investigations into the relevant phenomena. In this work, we present a recent implementation of chemical equilibrium, finite-rate chemistry, and thermochemical nonequilibrium capabilities into LASTRAC, an existing well-established boundary-layer stability analysis code. Verification against existing numerical results in the literature are presented. LASTRAC was previously able to address calorically perfect flows. By using solutions of the Parabolized Stability Equations (PSE) with chemical and thermal nonequilibrium, we are able to investigate the effects of chemical and thermal nonequilibrium on a variety of phenomena including stationary crossflow instability on a swept wing and 2nd mode instabilities over a wedge.

Kline, H. L.↗
NASA NTRSDate not supplied

NASA Langley Aerothermodynamics Laboratory: Hypersonic Testing Capabilities

A description of the NASA Langley Research Center’s Langley Aerothermodynamics Laboratory (LAL) will be presented in the paper, along with descriptions and details of the facility test techniques and recent upgrades. The LAL consists of three hypersonic blow-down wind tunnels covering Mach numbers of 6 and 10 and unit Reynolds number ranges of 0.5 to 8.3 million per foot as well as a 60-ft Vacuum Sphere Test Chamber. LAL facilities are used to study and define the aerodynamic performance and aeroheating characteristics of flight vehicle concepts. Data collected in the facilities have been used for design and optimization, anchoring computational predictions, generation of aerodynamic databases and design of Thermal Protection Systems. Over the years modifications and enhancements have been made to the facility hardware and instrumentation to increase efficiency, data quality, capabilities and reliability to better meet the programmatic requirements. Recent utilization information illustrates the need for the capabilities associated with these facilities. Recent test programs include the Space Shuttle Program, Crew Exploration Vehicle/Orion/Multi-Purpose Crew Vehicle, Hypersonic International Flight Research Experimentation (HIFiRE), Mars Science Laboratory, Hypersonic Inflatable Aerodynamic Decelerator System (HIADS) and X-51 among others and usage has been split between NASA, Commercial Crew, Department of Defense and private company programs. Plans for future improvements to the facility infrastructure and instrumentation will also be presented.

Karen Berger↗