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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 NTRS2016

Entry at Venus

This is lecture to be given at the IPPW 2016, as part of the 2 day course on Short Course on Destination Venus: Science, Technology and Mission Architectures. The attached presentation material is intended to be introduction to entry aspects of Venus in-situ robotic missions. The presentation introduces the audience to the aerodynamic and aerothermodynamic aspects as well as the loads, both aero and thermal, generated during entry. The course touches upon the system design aspects such as TPS design and both high and low ballistic coefficient entry system concepts that allow the science payload to be protected from the extreme entry environment and yet meet the mission objectives.

Thermal protection↗
NASA NTRS2017

Inviscid/Boundary-Layer Aeroheating Approach for Integrated Vehicle Design

A typical entry vehicle design depends on the synthesis of many essential subsystems, including thermal protection system (TPS), structures, payload, avionics, and propulsion, among others. The ability to incorporate aerothermodynamic considerations and TPS design into the early design phase is crucial, as both are closely coupled to the vehicle's aerodynamics, shape and mass. In the preliminary design stage, reasonably accurate results with rapid turn-representative entry envelope was explored. Initial results suggest that for Mach numbers ranging from 9-20, a few inviscid solutions could reasonably sup- port surface heating predictions at Mach numbers variation of +/-2, altitudes variation of +/-10 to 20 kft, and angle-of-attack variation of +/- 5. Agreement with Navier-Stokes solutions was generally found to be within 10-15% for Mach number and altitude, and 20% for angle of attack. A smaller angle-of-attack increment than the 5 deg around times for parametric studies and quickly evolving configurations are necessary to steer design decisions. This investigation considers the use of an unstructured 3D inviscid code in conjunction with an integral boundary-layer method; the former providing the flow field solution and the latter the surface heating. Sensitivity studies for Mach number, angle of attack, and altitude, examine the feasibility of using this approach to populate a representative entry flight envelope based on a limited set of inviscid solutions. Each inviscid solution is used to generate surface heating over the nearby trajectory space. A subset of a considered in this study is recommended. Results of the angle-of-attack sensitivity studies show that smaller increments may be needed for better heating predictions. The approach is well suited for application to conceptual multidisciplinary design and analysis studies where transient aeroheating environments are critical for vehicle TPS and thermal design. Concurrent prediction of aeroheating environments, coupled with the use of unstructured methods, is considered enabling for TPS material selection and design in conceptual studies where vehicle mission, shape, and entry strategies evolve rapidly.

Lee, Esther↗
NASA NTRS2017

Characterization of a Method for Inverse Heat Conduction Using Real and Simulated Thermocouple Data

It is often impractical to instrument the external surface of high-speed vehicles due to the aerothermodynamic heating. Temperatures can instead be measured internal to the structure using embedded thermocouples, and direct and inverse methods can then be used to estimate temperature and heat flux on the external surface. Two thermocouples embedded at different depths are required to solve direct and inverse problems, and filtering schemes are used to reduce noise in the measured data. Accuracy in the estimated surface temperature and heat flux is dependent on several factors. Factors include the thermocouple location through the thickness of a material, the sensitivity of the surface solution to the error in the specified location of the embedded thermocouples, and the sensitivity to the error in thermocouple data. The effect of these factors on solution accuracy is studied using the methodology discussed in the work of Pizzo, et. al.1 A numerical study is performed to determine if there is an optimal depth at which to embed one thermocouple through the thickness of a material assuming that a second thermocouple is installed on the back face. Solution accuracy will be discussed for a range of embedded thermocouple depths. Moreover, the sensitivity of the surface solution to (a) the error in the specified location of the embedded thermocouple and to (b) the error in the thermocouple data are quantified using numerical simulation, and the results are discussed.

Pizzo, Michelle E.↗
NASA NTRS2017

Modeling of a Turbofan Engine with Ice Crystal Ingestion in the NASA Propulsion System Laboratory

The main focus of this study is to apply a computational tool for the flow analysis of the turbine engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The PSL has been used to test a highly instrumented Honeywell ALF502R-5A (LF11) turbofan engine at simulated altitude operating conditions. Test data analysis with an engine cycle code and a compressor flow code was conducted to determine the values of key icing parameters, that can indicate the risk of ice accretion, which can lead to engine rollback (un-commanded loss of engine thrust). The full engine aerothermodynamic performance was modeled with the Honeywell Customer Deck specifically created for the ALF502R-5A engine. The mean-line compressor flow analysis code, which includes a code that models the state of the ice crystal, was used to model the air flow through the fan-core and low pressure compressor. The results of the compressor flow analyses included calculations of the ice-water flow rate to air flow rate ratio (IWAR), the local static wet bulb temperature, and the particle melt ratio throughout the flow field. It was found that the assumed particle size had a large effect on the particle melt ratio, and on the local wet bulb temperature. In this study the particle size was varied parametrically to produce a non-zero calculated melt ratio in the exit guide vane (EGV) region of the low pressure compressor (LPC) for the data points that experienced a growth of blockage there, and a subsequent engine called rollback (CRB). At data points where the engine experienced a CRB having the lowest wet bulb temperature of 492 degrees Rankine at the EGV trailing edge, the smallest particle size that produced a non-zero melt ratio (between 3 percent - 4 percent) was on the order of 1 micron. This value of melt ratio was utilized as the target for all other subsequent data points analyzed, while the particle size was varied from 1 micron - 9.5 microns to achieve the target melt ratio. For data points that did not experience a CRB which had static wet bulb temperatures in the EGV region below 492 degrees Rankine, a non-zero melt ratio could not be achieved even with a 1 micron ice particle size. The highest value of static wet bulb temperature for data points that experienced engine CRB was 498 degrees Rankine with a particle size of 9.5 microns. Based on this study of the LF11 engine test data, the range of static wet bulb temperature at the EGV exit for engine CRB was in the narrow range of 492 degrees Rankine - 498 degrees Rankine , while the minimum value of IWAR was 0.002. The rate of blockage growth due to ice accretion and boundary layer growth was estimated by scaling from a known blockage growth rate that was determined in a previous study. These results obtained from the LF11 engine analysis formed the basis of a unique “icing wedge.”

Turbo engine↗
NASA NTRS2015

Pilot Deployment of the LDSD Parachute via a Supersonic Ballute

The Low Density Supersonic Decelerator (LDSD) Project required the use of a pilot system due to the inability to mortar deploy its main supersonic parachute. A mortar deployed 4.4 m diameter supersonic ram-air ballute was selected as the pilot system for its high drag coefficient and stability relative to candidate supersonic parachutes at the targeted operational Mach number of 3. The ballute underwent a significant development program that included the development of a new liquid methanol-based pre-inflation system to assist the ballute inflation process. Both pneumatic and pyrotechnic mortar tests were conducted to verify orderly rigging deployment, bag strip, inflation aid activation, and proper mortar performance. The ballute was iteratively analyzed between fluid and structural analysis codes to obtain aerodynamic and aerothermodynamic estimates as well as estimates of the ballute's structural integrity and shape. The ballute was successfully flown in June 2014 at a Mach number of 2.73 as part of the first LDSD supersonic flight test and performed beyond expectations. Recovery of the ballute indicated that it did not exceed its structural or thermal capabilities. This flight set a historical precedent as it represented the largest ballute to have ever been successfully flown at this Mach number by a NASA entity.

Tanner, Christopher L.↗
NASA NTRS2017

Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) Sensor Suite

The Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite seeks to address the aerodynamic, aerothermodynamic, and thermal protection system (TPS) performance issues during atmospheric entry, descent, and landing of the Mars 2020 mission. Based on the highly successful instrumentation suite that flew on Mars Science Laboratory (MEDLI), the new sensor suite expands on the types of measurements and also seeks to answer questions not fully addressed by the previous mission. Sensor Package: MEDLI2 consists of 7 pressure transducers, 17 thermal plugs, 2 heat flux sensors, and one radiometer. The sensors are distributed across both the heatshield and backshell, unlike MEDLI (the first sensor suite), which was located solely on the heat-shield. The sensors will measure supersonic pressure on the forebody, a pressure measurement on the aftbody, near-surface and in-depth temperatures in the heatshield and backshell TPS materials, direct total heat flux on the aftbody, and direct radiative heating on the aftbody. Instrument Development: The supersonic pressure transducers, the direct heat flux sensors, and the radiometer all were tested during the development phase. The status of these sensors, including the piezo-resistive pressure sensors, will be presented. The current plans for qualification and calibration for all of the sensors will also be discussed. Post-Flight Data Analysis: Similar to MEDLI, the estimated flight trajectory will be reconstructed from the data. The aerodynamic parameters that will be reconstructed will be the axial force coefficient, freestream Mach number, base pressure, atmospheric density, and winds. The aerothermal quantities that will be determined are the heatshield and backshell aero-heating, turbulence transition across the heatshield, and TPS in-depth performance of PICA. By directly measuring the radiative and total heat fluxes on the back-shell, the convective portion of the heat flux will be estimated. The status of the current tools to perform the post-flight data analysis will be presented, along with plans for model improvements.

Hwang, Helen H.↗
NASA NTRS2017

Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) Sensor Suite

The Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite seeks to address the aerodynamic, aerothermodynamic, and thermal protection system (TPS) performance issues during atmospheric entry, descent, and landing of the Mars 2020 mission. Based on the highly successful instrumentation suite that flew on Mars Science Laboratory (MEDLI), the new sensor suite expands on the types of measurements and also seeks to answer questions not fully addressed by the previous mission. Sensor Package: MEDLI2 consists of 7 pressure transducers, 17 thermal plugs, 2 heat flux sensors, and one radiometer. The sensors are distributed across both the heatshield and backshell, unlike MEDLI (the first sensor suite), which was located solely on the heat-shield. The sensors will measure supersonic pressure on the forebody, a pressure measurement on the aftbody, near-surface and in-depth temperatures in the heatshield and backshell TPS materials, direct total heat flux on the aftbody, and direct radiative heating on the aftbody. Instrument Development: The supersonic pressure transducers, the direct heat flux sensors, and the radiometer all were tested during the development phase. The status of these sensors, including the piezo-resistive pressure sensors, will be presented. The current plans for qualification and calibration for all of the sensors will also be discussed. Post-Flight Data Analysis: Similar to MEDLI, the estimated flight trajectory will be reconstructed from the data. The aerodynamic parameters that will be reconstructed will be the axial force coefficient, freestream Mach number, base pressure, atmospheric density, and winds. The aerothermal quantities that will be determined are the heatshield and backshell aero-heating, turbulence transition across the heatshield, and TPS in-depth performance of PICA. By directly measuring the radiative and total heat fluxes on the back-shell, the convective portion of the heat flux will be estimated. The status of the current tools to perform the post-flight data analysis will be presented, along with plans for model improvements.

Hwang, Helen↗
NASA NTRS2018

A Reliability Comparison of Classical and Stochastic Thickness Margin Approaches to Address Material Property Uncertainties for the Orion Heat Shield

The Orion Thermal Protection System (TPS) margin process uses a root-sum-square approach with branches addressing trajectory, aerothermodynamics, and material response uncertainties in ablator thickness design. The material response branch applies a bondline temperature reduction between the Avcoat ablator and EA9394 adhesive by 60 C (108 F) from its peak allowed value of 260 C (500 F). This process is known as the Bond Line Temperature Material Margin (BTMM) and is intended to cover material property and performance uncertainties. The value of 60 C (108 F) is a constant, applied at any spacecraft body location and for any trajectory. By varying only material properties in a random (monte carlo) manner, the perl-based script mcCHAR is used to investigate the confidence interval provided by the BTMM. In particular, this study will look at various locations on the Orion heat shield forebody for a guided and an abort (ballistic) trajectory.

margins↗
NASA NTRS2018

A Reliability Comparison of Classical and Stochastic Thickness Margin Approaches to Address Material Property Uncertainties for the Orion Heat Shield

The Orion Thermal Protection System (TPS) margin process uses a root-sum-square approach with branches addressing trajectory, aerothermodynamics, and material response uncertainties in ablator thickness design. The material response branch applies a bond line temperature reduction between the Avcoat ablator and EA9394 adhesive by 60 C (108 F) from its peak allowed value of 260 C (500 F). This process is known as the Bond Line Temperature Material Margin (BTMM) and is intended to cover material property and performance uncertainties. The value of 60 C (108 F) is a constant, applied at any spacecraft body location and for any trajectory. By varying only material properties in a random (monte carlo) manner, the perl-based script mcCHAR is used to investigate the confidence interval provided by the BTMM. In particular, this study will look at various locations on the Orion heat shield forebody for a guided and an abort (ballistic) trajectory.

FIAT↗
NASA NTRS2018

Hybrid Reduced Order Model for N2-N2 Interactions for Application to Dissociation and Energy Transfer Processes

Recent work in the aerothermodynamics community has focused on the development of reduced-order models for thermo-chemical non-equilibrium which avoid the restrictive assumptions of multi-temperature models and the prohibitive cost associated with State-to-State (StS) models. In the present work, this is accomplished by lumping energy states together and assuming groups of states are roughly in equilibrium. As a result, the non-equilibrium behavior of a gas can be captured at a reduced computational cost from a full StS simulation. In this work, we present a hybrid grouping model for studying energy transfer and dissociation in a mixture of nitrogen molecules due to N2-N2 reactions. This is accomplished by making use of a grouping strategy informed by data from the N2-N StS kinetic data. However, due to the massive computational cost associated with constructing StS data for the N2-N2 system, the kinetic data for the hybrid grouping model are calculated using the quasi-classical trajectory (QCT) method by sampling states for trajectory within the groups. This general framework is called the Maximum-Entropy Quasi-Classical Trajectory (ME-QCT) method. The primary challenge associated with this method is that rates for reverse grouped reactions cannot be obtained through detailed balance at a group level, due to the variation of group internal temperatures. To construct the full model for N2-N2 grouped kinetics using the ME-QCT method, detailed balance is invoked at the microscopic level, allowing for the calculation of the full kinetic data from QCT. Results will be presented using the full ME-QCT model for the N2-N2 system in an isothermal and isochoric reactor simulation. In addition, simple CFD test cases for a one-dimensional standing shock and for a quasi-one-dimensional nozzle will be used for demonstration of the ME-QCT method. This method allows for the calculation of non-equilibrium behavior for the N2-N2 system without the prohibitive cost of a full StS simulation. Moreover, it enables the construction of a unified model for the dissociating and recombining non-equilibrium flows.

nonequilibrium modeling↗
NASA NTRS2018

Numerically Simulating an Expanding Continuum Jet into a Surrounding Non-Continuum Region

A numerical technique that simulates the interaction between a continuum jet and a free-molecular, rarefied or transitional flow field is presented and applied. The technique is based on selecting a boundary between the expanding jet plume and interacting flow field defined by the Bird breakdown parameter, P. By properly choosing the boundary, the jet plume computed by computational fluid dynamics (CFD) is uncoupled from the interaction region that is simulated by direct simulation Monte Carlo (DSMC). The mechanics of uncoupling the continuum and non-continuum regions is discussed first. Also, a discussion of expanding continuum plume breakdown and the formulation of the Bird breakdown parameter is presented. To show the usefulness, examples of applying the CFD-DSMC technique are given. Many of the examples were taken from projects worked by the staff of the Aerothermodynamics Branch at the NASA Langley Research Center. These examples show a variety of applications of the technique and allow those examples not previously documented to be formally presented.

Jet interaction↗
NASA NTRS2018

Imaging for Hypersonic Experimental Aeroheating Testing (IHEAT) Version 4.0: User Manual

The IHEAT v4.0 software is a data reduction code for global thermography data acquired in the NASA Langley Aerothermodynamics Laboratory (LAL) hypersonic wind tunnels. IHEAT uses red and green color-intensity data from two-dimensional images of wind tunnel models to compute temperatures and heat-transfer rates using a semi-infinite, one-dimensional heat transfer approximation at each image pixel. Multiple automated tools in IHEAT v4.0 decrease the time required to reduce the data from a phosphor thermography wind tunnel run. Data at one or all of the image pixel locations can be exported to computer files for further analysis. The prior version of IHEAT, v3.2, was written in PV-WAVE® (now owned by Rogue Wave® Software) in 1994 and was limited in functionality to fit within the memory constraints of the available computers at the time. IHEAT v4.0 is written in MATLAB® by MathWorks® and contains several new features that leverage the increase in available memory of the current computers. A Piecewise tool permits the user to extract data along a segmented line cut that can follow interesting features in the image better than the single, straight line cuts that were possible with the legacy Length and Profile tools. The new Load Run and Batch tools facilitate batch processing by loading in all of the input files and images for a run at the same time. Load Run permits the user to process the available run images manually, while Batch automatically saves heat transfer data from all of the images based on the analysis previously performed on a single frame. IHEAT v4.0 also can automatically calculate the temporal collapse of reference line cuts from the time history heating data for a run to indicate the appropriate frame to reduce for each run. The IHEAT v4.0 source code was compiled into a standalone executable file that can be accessed remotely from several computers with different operating systems, simultaneously. The software is run through the MATLAB® Compiler Runtime engine, and therefore, IHEAT does not require a software license to run. Any software commands executed in the IHEAT v4.0 code will not affect other similar applications running on the same machine. Similarly, changes to the parent software do not affect a compiled code. These features of IHEAT v4.0 are improvements over the legacy v3.2 code, which required regular maintenance to avoid losing functionality as the PVWAVE ® programming language was upgraded.

Mason, Michelle L.↗
NASA NTRS2018

Studies in Support of Venus Aerocapture Utilizing Drag Modulation

Aerocapture has been extensively studied and these studies have shown the benefit for planetary exploration missions. While the traditional approach to aerocapture with lifting configurations and lift-guided modulations have been assessed to be technologically feasible, aerocapture using purely drag modulation was proposed and studied by Prof. Braun and his students. These studies show that if one can assess the feasibility of aerocapture using drag modulation at Venus, and develop tall pole technologies needed at Venus, then this concept is much easier to execute at all other relevant destinations. Based on the above finding, partnered proposals were submitted by Adam Nelessen at JPL and Ethiraj Venkatapathy at Ames in collaboration with Prof. Braun at the University of Colorado, Boulder (UCB). Under this partnership, Ames Research Center (ARC) is working to address some of the key entry technology challenges associated with drag modulation aerocapture at Venus. Drag modulation aerocapture is a simple, scalable, and likely cost-effective way to enhance planetary science missions. The approach envisioned is to design a small spacecraft, that would most likely be a secondary payload, with a removable drag skirt. The vehicle would enter the atmosphere at Venus with a low ballistic coefficient, decelerate rapidly, drop the skirt resulting in a smaller vehicle with a higher ballistic coefficient which would skip out of the atmosphere and enter into a desired orbit. ARC's role in this collaboration is multifold. First of which is to perform design studies on various pre- and post-jettison geometries utilizing a 3-DOF trajectory code to determine the aerodynamics and aerothermodynamics of the vehicles and evaluate viable thermal protection material system designs. Once these design studies are complete, Ames will then perform higher fidelity CFD and TPS sizing to further design the vehicles. Second, the multi-body separation dynamics of the drag modulation event will be explored using both CFD simulations (CART3-D and US3D) as well as possible ballistic range testing. ARC's tools and expertise have been used to assess and advise on the selection of the separating configuration. In addition to the preliminary evaluation, ARC will provide tools and expertise to UCB team members to further assess aerodynamic interactions between the separating bodies and provide guidance as to the feasibility of stable transition.

Beck, Robin↗
NASA NTRS2019

InSight's Reconstructed Aerothermal Environments

The InSight Mars Lander successfully landed on the surface on November 26, 2018. This poster will describe the methodologies and margins used in developing the aerothermal environments for design of the thermal protection systems (TPS), as well as a prediction of as-flown environments based on the best estimated trajectory. The InSight mission spacecraft design approach included the effects of radiant heat flux to the aft body from the wake for the first time on a US Mars Mission, due to overwhelming evidence in ground testing for the European ExoMars mission (2009/2010) [1] and 2010 tests in the Electric Arc Shock Tube (EAST) facility [2]. The radiant energy on an aftbody was also recently confirmed via measurement on the Schiaparelli mission [3]. In addition, the InSight mission expected to enter the Mars atmosphere during the dust storm season, so the heatshield TPS was designed to accommodate the extra recession due to the potential dust impact. This poster will compare the predicted aerothermal environments using the reconstructed best estimated trajectory to the design environments. Design Approach: The InSight spacecraft was planned to be a near-design-to-print copy of the Phoenix spacecraft. The determination of the heatshield TPS requirements was approached as if it was a new design due to the new requirement of flying through a dust storm. The baseline for aftbody was build-to-print, and all analyses focused on ensuring adequate margin. This proved to be a challenge because the Phoenix aftbody was designed to withstand only convective heating and the InSight aftbody was evaluated for both convective and radiative heating. Aerothermal environments were predicted using the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) and the Data Parallel Line Relaxation (DPLR) CFD codes, and the Nonequilibrium Radiative Transport and Spectra Program (NEQAIR) utilizing bounding design trajectories derived from Monte Carlo analyses from the Program to Optimize Simulated Trajectories II (POST2). In all cases, super-catalytic flowfields were assigned to ensure the most conservative heating results. Two trajectories were evaluated: 1) the trajectory with the maximum heat flux was utilized to determine the flowfield characteristics and the viability of the selection of TPS materials; and 2) the trajectory with the maximum heat load was used to determine the required thicknesses of the TPS materials. Evaluation of the MEDLI data [4], along with ground test data [5] led to the determination of whether or not the flow would transition from laminar to turbulent on the heatshield, which also determined the TPS sizing location for the heatshield. Aerothermal margins were added for the convective heating and developed for the radiative heating. TPS material sizing was determined with the Reaction Kinetic Ablation Program (REKAP) and the Fully Implicit Ablation and Thermal Analysis program (FIAT) using a three-branched approach to account for aerothermal, material response, and material properties uncertainties. In addition, the heatshield recession was augmented by an analysis of the effect of entry through a potential dusty atmosphere using a methodology developed in References [6] and [7]. These analyses resulted in an increase to the Phoenix heatshield TPS thickness. Reconstruction Efforts: Once the best estimated trajectory is reconstructed by the team, the LAURA/HARA (High-Temperature Aerothermo-dynamic Radiation model) and DPLR/NEQAIR code pairs will be used to predict the as-flown aerothermal conditions. In these runs, fully-catalytic flowfields will be assigned because it is a more physically accurate description of the chemistry in the flow. Once again, determination of the onset of turbulence on the heatshield will be evaluated. The as-flown aerothermal environments will then be compared to the design environments.

Beck, R. A.↗
NASA NTRS2018

Comparison of Several Global Mixing Performance Metrics for High-Speed Fuel Injectors

To experimentally assess and compare the mixing performance of high-speed fuel injectors for scramjet engines, quantitative global metrics are needed. The one-dimensional metric most commonly used to assess the degree of mixing completeness at a given downstream station is the mixing efficiency parameter. The experimental determination of the mixing efficiency parameter requires measurement of the spatial distributions of both the fuel mass fraction and the mass flux. Standard in-stream gas sampling techniques can be used to measure the fuel mass fraction distribution, however the mass flux distribution is not easily determined experimentally because it requires the measurement of three independent aerothermodynamic variables in addition to the mixture composition. For this reason, several metrics that can be calculated from the fuel distribution alone are commonly used to assess mixing performance. Because these other metrics do not provide a mass flux-weighted measure of the local degree of mixing completeness, they may not correlate well with the mixing efficiency parameter. Therefore, if the substitute metrics are to be used to compare the mixing performance of candidate fuel injector concepts, it is important to understand their relationships to the mixing efficiency parameter in a representative scramjet combustor flowfield. This work investigates the relationships between the mixing efficiency parameter and several substitute metrics that are able to be measured with the current experimental setup of the Enhanced Injection and Mixing Project at the NASA Langley Research Center for baseline strut and ramp injectors. The results of these comparisons have revealed that it is possible to glean different (i.e., incorrect) conclusions about which injector is the better mixer when the substitute mixing performance metrics are used instead of the mixing efficiency parameter, thereby highlighting the importance of mass flux-weighted mixing performance metrics.

Ground, Cody R.↗
NASA NTRS2017

Implementation of a Body Force Model in OVERFLOW for Propulsor Simulations

We present an implementation of a propulsor model based on body force method into the OVERFLOW computational fluid dynamics code to model turbofan engines and propulsors of similar type. The model estimates the forces imparted on the fluid by the blade camber surfaces as a body force source terms in the momentum and energy equations over grids that represent the rotor and stator of a fan stage. We tested the implementation on three cases: 1) Source Diagnostics Test (SDT) fan with R4 rotors, which has extensive test data on aerodynamic performance and rotor wake surveys, 2) A stand-alone Aeronaut TF8000 propulsor which is an off-the-shelf propulsor used on model aircraft and 3) The D8 aircraft model with TF8000 propulsors placed in a wind tunnel. Despite missing some of the features in the velocity profiles near the endwalls (i.e. hub and casing surfaces), the OVERFLOW simulations with body force model predicted area-averaged flow speed and total pressure rise through the SDT propulsor within a few percent of the LDV (Laser Doppler Velocimetry) measurements. In the case of TF8000 propulsor on the D8 airframe, the model under-predicted mechanical power coefficient by several percent of the wind tunnel test results when the horizontal force balance condition over the airframe is targeted by tuning rotor speed. By investigating the upstream influence of the rotor swirl, it was found out that the induced swirl velocity effects upstream were relatively small and they rapidly vanished before reaching one fan diameter upstream of the fan face. The body force model provided insights on aerothermodynamics and aeromechanics of boundary layer ingesting propulsor; these insights could not be obtained by using the uniform pressure jump model.

Propulsor↗