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NASA NTRS1995

Validation of the SINDA/FLUINT code using several analytical solutions

The Systems Improved Numerical Differencing Analyzer and Fluid Integrator (SINDA/FLUINT) code has often been used to determine the transient and steady-state response of various thermal and fluid flow networks. While this code is an often used design and analysis tool, the validation of this program has been limited to a few simple studies. For the current study, the SINDA/FLUINT code was compared to four different analytical solutions. The thermal analyzer portion of the code (conduction and radiative heat transfer, SINDA portion) was first compared to two separate solutions. The first comparison examined a semi-infinite slab with a periodic surface temperature boundary condition. Next, a small, uniform temperature object (lumped capacitance) was allowed to radiate to a fixed temperature sink. The fluid portion of the code (FLUINT) was also compared to two different analytical solutions. The first study examined a tank filling process by an ideal gas in which there is both control volume work and heat transfer. The final comparison considered the flow in a pipe joining two infinite reservoirs of pressure. The results of all these studies showed that for the situations examined here, the SINDA/FLUINT code was able to match the results of the analytical solutions.

Keller, John R.↗

NASA NTRS2008

Validation of Ray Tracing Code Refraction Effects

NASA's current predictive capabilities using the ray tracing program (RTP) are validated using helicopter noise data taken at Eglin Air Force Base in 2007. By including refractive propagation effects due to wind and temperature, the ray tracing code is able to explain large variations in the data observed during the flight test.

Heath, Stephanie L.↗

NASA NTRS2002

High-Pressure Gaseous Burner (HPGB) Facility Completed for Quantitative Laser Diagnostics Calibration

A gas-fueled high-pressure combustion facility with optical access, which was developed over the last 2 years, has just been completed. The High Pressure Gaseous Burner (HPGB) rig at the NASA Glenn Research Center can operate at sustained pressures up to 60 atm with a variety of gaseous fuels and liquid jet fuel. The facility is unique as it is the only continuous-flow, hydrogen-capable, 60-atm rig in the world with optical access. It will provide researchers with new insights into flame conditions that simulate the environment inside the ultra-high-pressure-ratio combustion chambers of tomorrow's advanced aircraft engines. The facility provides optical access to the flame zone, enabling the calibration of nonintrusive optical diagnostics to measure chemical species and temperature. The data from the HPGB rig enables the validation of numerical codes that simulate gas turbine combustors, such as the National Combustor Code (NCC). The validation of such numerical codes is often best achieved with nonintrusive optical diagnostic techniques that meet these goals: information-rich (multispecies) and quantitative while providing good spatial and time resolution. Achieving these goals is a challenge for most nonintrusive optical diagnostic techniques. Raman scattering is a technique that meets these challenges. Raman scattering occurs when intense laser light interacts with molecules to radiate light at a shifted wavelength (known as the Raman shift). This shift in wavelength is unique to each chemical species and provides a "fingerprint" of the different species present. The facility will first be used to gather a comprehensive data base of laser Raman spectra at high pressures. These calibration data will then be used to quantify future laser Raman measurements of chemical species concentration and temperature in this facility and other facilities that use Raman scattering.

Nguyen, Quang-Viet↗

NASA NTRS2000

Flow Duct Data for Validation of Acoustic Liner Codes for Impedance Eduction

The objective of the study reported here was to acquire acoustic and flow data with hard and lined duct wall duct sections for validation of a liner prediction code being developed at NASA LaRC. Both the mean flowfield and acoustic flowfields were determined in a cross-plane of the rectangular duct. A flow duct facility with acoustic drivers connected to a rectangular (4.7 x 2.0 inch) source section and a linear acoustic liner mounted downstream of the source section was used in this study. The liner section was designed to allow liner materials to be placed on all 4 walls of the duct. The test liner was of the locally-reacting type and was made from a ceramic material. The material, consisting of a tubular structure, was provided by NASA LaRC. The liner was approximately 8.89 cm (3.5 inches) thick. For the current study, only the two "short" sides of the duct were lined with liner material. The other two sides were hard walls. Two especially built instrumentation sections were attached on either sides of the liner section to allow acoustic and flow measurements to be made upstream and downstream of the liner. The two instrumentation duct sections were built to allow measurement of acoustic and flow properties at planes perpendicular to flow upstream and downstream of the liner section. The instrumentation section was also designed to provide a streamwise gradient in acoustic (complex) pressure from which the acoustic particle velocity, needed for the model validation, can be computed. Flow measurements included pressure, temperature, and velocity profiles upstream of the liner section. The in-flow sound pressure levels and phases were obtained with a microphone probe equipped with a nose cone in two cross planes upstream of the liner and two cross plane downstream of the liner. In addition to the acoustic measurements at the cross planes. axial centerline acoustic data was acquired using an axially traversing microphone probe which was traversed from a location upstream of the liner to some distance downstream of the liner. All probes used here had to be calibrated with respect to a standard microphone equipped with a nose cone to allow for the effects of flow.

Ahuja, K. K.↗

NASA NTRS1992

Activities of the NASA/Marshall Space Flight Center pump stage technology team

In order to advance rocket propulsion technology, the Consortium for Computational Fluid Dynamics (CFD) Application in Propulsion Technology has been formed at Marshall Space Flight Center (MSFC). The Consortium consists of three Teams: the turbine stage team, the pump stage team (PST), and the combustion devices team. The PST has formulated and is implementing a plan for pump technology development whose end product will be validated CFD codes suitable for application to pump components, test data suitable for validating CFD codes, and advanced pump components optimized using CFD codes. The PST's work during the fall of 1991 and the winter and spring of 1992 is discussed in this paper. This work is highlighted by CFD analyses of an advanced impeller design and collection of laser two-focus velocimeter data for the Space Shuttle Main Engine High Pressure Fuel Pump impeller.

Garcia, R.↗

NASA NTRS1986

Composite load spectra for select space propulsion structural components

A multiyear program is performed with the objective to develop generic load models with multiple levels of progressive sophistication to simulate the composite (combined) load spectra that are induced in space propulsion system components, representative of Space Shuttle Main Engines (SSME), such as transfer ducts, turbine blades, and liquid oxygen (LOX) posts. Progress of the first year's effort includes completion of a sufficient portion of each task -- probabilistic models, code development, validation, and an initial operational code. This code has from its inception an expert system philosophy that could be added to throughout the program and in the future. The initial operational code is only applicable to turbine blade type loadings. The probabilistic model included in the operational code has fitting routines for loads that utilize a modified Discrete Probabilistic Distribution termed RASCAL, a barrier crossing method and a Monte Carlo method. An initial load model was developed by Battelle that is currently used for the slowly varying duty cycle type loading. The intent is to use the model and related codes essentially in the current form for all loads that are based on measured or calculated data that have followed a slowly varying profile.

Newell, J. F.↗

NASA NTRS1995

New higher-order Godunov code for modelling performance of two-stage light gas guns

A new quasi-one-dimensional Godunov code for modeling two-stage light gas guns is described. The code is third-order accurate in space and second-order accurate in time. A very accurate Riemann solver is used. Friction and heat transfer to the tube wall for gases and dense media are modeled and a simple nonequilibrium turbulence model is used for gas flows. The code also models gunpowder burn in the first-stage breech. Realistic equations of state (EOS) are used for all media. The code was validated against exact solutions of Riemann's shock-tube problem, impact of dense media slabs at velocities up to 20 km/sec, flow through a supersonic convergent-divergent nozzle and burning of gunpowder in a closed bomb. Excellent validation results were obtained. The code was then used to predict the performance of two light gas guns (1.5 in. and 0.28 in.) in service at the Ames Research Center. The code predictions were compared with measured pressure histories in the powder chamber and pump tube and with measured piston and projectile velocities. Very good agreement between computational fluid dynamics (CFD) predictions and measurements was obtained. Actual powder-burn rates in the gun were found to be considerably higher (60-90 percent) than predicted by the manufacturer and the behavior of the piston upon yielding appears to differ greatly from that suggested by low-strain rate tests.

Bogdanoff, D. W.↗

NASA NTRS1998

Application of Finite Element Method to Analyze Inflatable Waveguide Structures

A Finite Element Method (FEM) is presented to determine propagation characteristics of deformed inflatable rectangular waveguide. Various deformations that might be present in an inflatable waveguide are analyzed using the FEM. The FEM procedure and the code developed here are so general that they can be used for any other deformations that are not considered in this report. The code is validated by applying the present code to rectangular waveguide without any deformations and comparing the numerical results with earlier published results.

Deshpande, M. D.↗

NASA NTRS1994

Validation of the RPLUS3D Code for Supersonic Inlet Applications Involving Three-Dimensional Shock Wave-Boundary Layer Interactions

A three-dimensional computational fluid dynamics code, RPLUS3D, which was developed for the reactive propulsive flows of ramjets and scramjets, was validated for glancing shock wave-boundary layer interactions. Both laminar and turbulent flows were studied. A supersonic flow over a wedge mounted on a flat plate was numerically simulated. For the laminar case, the static pressure distribution, velocity vectors, and particle traces on the flat plate were obtained. For turbulent flow, both the Baldwin-Lomax and Chien two-equation turbulent models were used. The static pressure distributions, pitot pressure, and yaw angle profiles were computed. In addition, the velocity vectors and particle traces on the flat plate were also obtained from the computed solution. Overall, the computed results for both laminar and turbulent cases compared very well with the experimentally obtained data.

Kapoor, Kamlesh↗

NASA NTRS1998

Euler Calculations at Off-Design Conditions for an Inlet of Inward Turning RBCC-SSTO Vehicle

The inviscid performance of an inward turning inlet design is calculated computationally for the first time. Hypersonic vehicle designs based on the inward turning inlets have been shown analytically to have increased effective specific impulse and lower heat load than comparably designed vehicles with two-dimensional inlets. The inward turning inlets are designed inversely from inviscid stream surfaces of known flow fields. The computational study is performed on a Mach 12 inlet design to validate the performance predicted by the design code (HAVDAC) and calculate its off-design Mach number performance. The three-dimensional Euler equations are solved for Mach 4, 8, and 12 using a software package called SAM, which consists of an unstructured mesh generator (SAMmesh), a three-dimensional unstructured mesh flow solver (SAMcfd), and a CAD-based software (SAMcad). The computed momentum averaged inlet throat pressure is within 6% of the design inlet throat pressure. The mass-flux at the inlet throat is also within 7 % of the value predicted by the design code thereby validating the accuracy of the design code. The off-design Mach number results show that flow spillage is minimal, and the variation in the mass capture ratio with Mach number is comparable to an ideal 2-D inlet. The results from the inviscid flow calculations of a Mach 12 inward turning inlet indicate that the inlet design has very good on and off-design performance which makes it a promising design candidate for future air-breathing hypersonic vehicles.

Takashima, N.↗

NASA NTRS2000

THRSTER: A Three-Stream Ejector Ramjet Analysis and Design Tool

An engineering tool for analyzing ejectors in rocket based combined cycle (RBCC) engines has been developed. A key technology for multi-cycle RBCC propulsion systems is the ejector which functions as the compression stage of the ejector ramjet cycle. The THRee STream Ejector Ramjet analysis tool was developed to analyze the complex aerothermodynamic and combustion processes that occur in this device. The formulated model consists of three quasi-one-dimensional streams, one each for the ejector primary flow, the secondary flow, and the mixed region. The model space marches through the mixer, combustor, and nozzle to evaluate the solution along the engine. In its present form, the model is intended for an analysis mode in which the diffusion rates of the primary and secondary into the mixed stream are stipulated. The model offers the ability to analyze the highly two-dimensional ejector flowfield while still benefits from the simplicity and speed of an engineering tool. To validate the developed code, wall static pressure measurements from the Penn-State and NASA-ART RBCC experiments were used to compare with the results generated by the code. The calculated solutions were generally found to have satisfactory agreement with the pressure measurements along the engines, although further modeling effort may be required when a strong shock train is formed at the rocket exhaust. The range of parameters in which the code would generate valid results are presented and discussed.

Chue, R. S.↗

NASA NTRS2000

THRSTER: A THRee-STream Ejector Ramjet Analysis and Design Tool

Chue, R. S.↗

NASA NTRS2001

New Gear Transmission Error Measurement System Designed

The prime source of vibration and noise in a gear system is the transmission error between the meshing gears. Transmission error is caused by manufacturing inaccuracy, mounting errors, and elastic deflections under load. Gear designers often attempt to compensate for transmission error by modifying gear teeth. This is done traditionally by a rough "rule of thumb" or more recently under the guidance of an analytical code. In order for a designer to have confidence in a code, the code must be validated through experiment. NASA Glenn Research Center contracted with the Design Unit of the University of Newcastle in England for a system to measure the transmission error of spur and helical test gears in the NASA Gear Noise Rig. The new system measures transmission error optically by means of light beams directed by lenses and prisms through gratings mounted on the gear shafts. The amount of light that passes through both gratings is directly proportional to the transmission error of the gears. A photodetector circuit converts the light to an analog electrical signal. To increase accuracy and reduce "noise" due to transverse vibration, there are parallel light paths at the top and bottom of the gears. The two signals are subtracted via differential amplifiers in the electronics package. The output of the system is 40 mV/mm, giving a resolution in the time domain of better than 0.1 mm, and discrimination in the frequency domain of better than 0.01 mm. The new system will be used to validate gear analytical codes and to investigate mechanisms that produce vibration and noise in parallel axis gears.

Oswald, Fred B.↗

NASA NTRS2005

Laminar Heating Validation of the OVERFLOW Code

OVERFLOW, a structured finite difference code, was applied to the solution of hypersonic laminar flow over several configurations assuming perfect gas chemistry. By testing OVERFLOW's capabilities over several configurations encompassing a variety of flow physics a validated laminar heating was produced. Configurations tested were a flat plate at 0 degrees incidence, a sphere, a compression ramp, and the X-38 re-entry vehicle. This variety of test cases shows the ability of the code to predict boundary layer flow, stagnation heating, laminar separation with re-attachment heating, and complex flow over a three-dimensional body. In addition, grid resolutions studies were done to give recommendations for the correct number of off-body points to be applied to generic problems and for wall-spacing values to capture heat transfer and skin friction. Numerical results show good comparison to the test data for all the configurations.

Lillard, Randolph P.↗

NASA NTRS2015

Jet-Surface Interaction: High Aspect Ratio Nozzle Test, Nozzle Design and Preliminary Data

The Jet-Surface Interaction High Aspect Ratio (JSI-HAR) nozzle test is part of an ongoing effort to measure and predict the noise created when an aircraft engine exhausts close to an airframe surface. The JSI-HAR test is focused on parameters derived from the Turbo-electric Distributed Propulsion (TeDP) concept aircraft which include a high-aspect ratio mailslot exhaust nozzle, internal septa, and an aft deck. The size and mass flow rate limits of the test rig also limited the test nozzle to a 16:1 aspect ratio, half the approximately 32:1 on the TeDP concept. Also, unlike the aircraft, the test nozzle must transition from a single round duct on the High Flow Jet Exit Rig, located in the AeroAcoustic Propulsion Laboratory at the NASA Glenn Research Center, to the rectangular shape at the nozzle exit. A parametric nozzle design method was developed to design three low noise round-to-rectangular transitions, with 8:1, 12:1, and 16: aspect ratios, that minimizes flow separations and shocks while providing a flat flow profile at the nozzle exit. These designs validated using the WIND-US CFD code. A preliminary analysis of the test data shows that the actual flow profile is close to that predicted and that the noise results appear consistent with data from previous, smaller scale, tests. The JSI-HAR test is ongoing through October 2015. The results shown in the presentation are intended to provide an overview of the test and a first look at the preliminary results.

Aeroacoustics↗

NASA NTRS2019

NASA's Role in Gas Turbine Technology Development: Accelerating Technical Progress via Collaboration Between Academia, Industry, and Government Agencies

Given the maturity of the gas turbine engine since its invention and also considering the limited and flattened level of resources expected to be allocated for NASA aeronautics research and development, we ask the question are NASA technology investments still needed to enable future turbine engine-based propulsion systems? If so, what is NASA’s unique role to justify NASA’s investment? To address this topic, we will first review the accomplishments and the impact that NASA Glenn Research Center has made on turbine engine technologies over the last 78 years. Specifically, this paper discusses NASA’s role and contributions to turbine engine development, specific to both 1) NASA’s role in conducting experiments to understand flow physics and provide relevant benchmark validation experiments for Computational Fluid Dynamics (CFD) code development, validation, and assessment; and 2) the impact of technologies resulting from NASA collaborations with industry, academia, and other government agencies. Note that the scope of the discussion is limited to the NASA technology contributions with which the author was intimately associated, and does not represent the entirety of the NASA contributions to turbine engine technology. The specific research, development, and demonstrations discussed herein were selected to both 1) provide a comprehensive review and reference list of the technology and its impact, and 2) identify NASA’s unique role and highlight how NASA’s involvement resulted in additional benefit to the gas turbine engine community. Secondly, we will discuss current NASA collaborations that are in progress and provide a status of the results. Finally, we discuss the challenges anticipated for future turbine engine-based propulsion systems for civil aviation and identify potential opportunities for collaboration where NASA involvement would be beneficial. Ultimately, the gas turbine engine community will decide if NASA involvement is needed to contribute to the development of the design and analysis tools, databases, and technology demonstration programs to meet these challenges for future turbine engine-based propulsion systems.

Suder, Kenneth L.↗

NASA NTRS2006

NASA Radiation Protection Research for Exploration Missions

The HZETRN code was used in recent trade studies for renewed lunar exploration and currently used in engineering development of the next generation of space vehicles, habitats, and EVA equipment. A new version of the HZETRN code capable of simulating high charge and energy (HZE) ions, light-ions and neutrons with either laboratory or space boundary conditions with enhanced neutron and light-ion propagation is under development. Atomic and nuclear model requirements to support that development will be discussed. Such engineering design codes require establishing validation processes using laboratory ion beams and space flight measurements in realistic geometries. We discuss limitations of code validation due to the currently available data and recommend priorities for new data sets.

Wilson, John W.↗

NASA NTRS1991

Infrared imaging - A validation technique for computational fluid dynamics codes used in STOVL applications

The need for a validation technique for computational fluid dynamics (CFD) codes in STOVL applications has led to research efforts to apply infrared thermal imaging techniques to visualize gaseous flow fields. Specifically, a heated, free-jet test facility was constructed. The gaseous flow field of the jet exhaust was characterized using an infrared imaging technique in the 2 to 5.6 micron wavelength band as well as conventional pitot tube and thermocouple methods. These infrared images are compared to computer-generated images using the equations of radiative exchange based on the temperature distribution in the jet exhaust measured with the thermocouple traverses. Temperature and velocity measurement techniques, infrared imaging, and the computer model of the infrared imaging technique are presented and discussed. From the study, it is concluded that infrared imaging techniques coupled with the radiative exchange equations applied to CFD models are a valid method to qualitatively verify CFD codes used in STOVL applications.

Hardman, R. R.↗