Engineering topics

Hirt, Stefanie M.

Publications and source records attributed to Hirt, Stefanie M..

18 records

NASA NTRS2018

Data Analysis Techniques for Fan Performance in Highly-Distorted Flows from Boundary Layer Ingesting Inlets

The design of a unique distortion-tolerant fan for a high-bypass ratio boundary-layer ingesting propulsion system has been completed and a rig constructed and tested in the NASA Glenn 8’x6’ wind tunnel. Processing the data from the experiment presented some interesting challenges because of the complexity of the experimental setup and the flow through the test rig. The experiment was run in three phases, each of which employed a unique complement of inlet throat and fan face instrumentation to avoid the blockage that would have resulted from simultaneously installing all of the rakes. The measurement from the individual test points were subsequently combined to compute the overall stage performance. A CFD model of the experiment was used to gain understanding of the flow field and to test some of the techniques proposed for interpolating and extrapolating the measurements into regions where measurements were not made. This capability became extremely useful when it was discovered that there was an unexpected total temperature distortion in the tunnel. The CFD model was modified by inserting a total temperature profile at the upstream boundary that mimicked the measured distortion where measurements were available and that CFD solution was used to investigate methods to infer the complete total temperature field at the fan face.

Hardin, Larry W.↗

NASA NTRS2018

Performance Calculations for a Boundary-Layer-Ingesting Fan Stage from Sparse Measurements

A test of the Boundary Layer Ingesting-Inlet / Distortion-Tolerant Fan was completed in NASA Glenn's 8-Foot by 6-Foot supersonic wind tunnel. Inlet and fan performance were measured by surveys using a set of rotating rake arrays upstream and downstream of the fan stage. Surveys were conducted along the 100 percent speed line and a constant exit corrected flow line passing through the aerodynamic design point. These surveys represented only a small fraction of the data collected during the test. For other operating points, data was recorded as snapshots without rotating the rakes which resulted in a sparser set of recorded data. This paper will discuss analysis of these additional, lower measurement density data points to expand our coverage of the fan map. Several techniques will be used to supplement the snapshot data at test conditions where survey data also exists. The supplemented snapshot data will be compared with survey results to assess the quality of the approach. Effective methods will be used to analyze the data set for which only snapshots exist.

fan efficiency↗

NASA NTRS2017

Development of a Rotating Rake Array for Boundary-Layer-Ingesting Fan-Stage Measurements

The recent Boundary-Layer-Ingesting Inlet/Distortion Tolerant Fan wind tunnel experiment at NASA Glenn Research Center's 8- by 6-foot Supersonic Wind Tunnel (SWT) examined the performance of a novel inlet and fan stage that was designed to ingest the vehicle boundary layer in order to take advantage of a predicted overall propulsive efficiency benefit. A key piece of the experiment's instrumentation was a pair of rotating rake arrays located upstream and downstream of the fan stage. This paper examines the development of these rake arrays. Pre-test numerical solutions were sampled to determine placement and spacing for rake pressure and temperature probes. The effects of probe spacing and survey density on the repeatability of survey measurements was examined. These data were then used to estimate measurement uncertainty for the adiabatic efficiency.

turbomachinery↗

NASA NTRS2017

Development of a Flow Field for Testing a Boundary-Layer-Ingesting Propulsor

The test section of the 8- by 6-Foot Supersonic Wind Tunnel at NASA Glenn Research Center was modified to produce the test conditions for a boundary-layer-ingesting propulsor. A test was conducted to measure the flow properties in the modified test section before the propulsor was installed. Measured boundary layer and freestream conditions were compared to results from computational fluid dynamics simulations of the external surface for the reference vehicle. Testing showed that the desired freestream conditions and boundary layer thickness could be achieved; however, some non-uniformity of the freestream conditions, particularly the total temperature, were observed.

embedded propulsion↗

NASA NTRS2017

Development of a Rotating Rake Array for Boundary-Layer-Ingesting Fan-Stage Measurements

The recent Boundary-Layer-Ingesting Inlet/Distortion Tolerant Fan wind tunnel experiment at NASA Glenn Research Center's 8-foot by 6-foot supersonic wind tunnel examined the performance of a novel inlet and fan stage that was designed to ingest the vehicle boundary layer in order to take advantage of a predicted overall propulsive efficiency benefit. A key piece of the experiment's instrumentation was a pair of rotating rake arrays located upstream and downstream of the fan stage. This paper examines the development of these rake arrays. Pre-test numerical solutions were sampled to determine placement and spacing for rake pressure and temperature probes. The effects of probe spacing and survey density on the repeatability of survey measurements was examined. These data were then used to estimate measurement uncertainty for the adiabatic efficiency.

Wolter, John D.↗

NASA NTRS2015

Experimental Study of Fillets to Reduce Corner Effects in an Oblique Shock-Wave/Boundary Layer Interaction

A test was conducted in the 15 cm x 15 cm supersonic wind tunnel at NASA Glenn Research Center that focused on corner effects of an oblique shock-wave/boundary-layer interaction. In an attempt to control the interaction in the corner region, eight corner fillet configurations were tested. Three parameters were considered for the fillet configurations: the radius, the fillet length, and the taper length from the square corner to the fillet radius. Fillets effectively reduced the boundary-layer thickness in the corner; however, there was an associated penalty in the form of increased boundary-layer thickness at the tunnel centerline. Larger fillet radii caused greater reductions in boundary-layer thickness along the corner bisector. To a lesser, but measureable, extent, shorter fillet lengths resulted in thinner corner boundary layers. Overall, of the configurations tested, the largest radius resulted in the best combination of control in the corner, evidenced by a reduction in boundary-layer thickness, coupled with minimal impacts at the tunnel centerline.

Fillets↗

NASA NTRS2014

Performance of a Supersonic Over-Wing Inlet with Application to a Low-Sonic-Boom Aircraft

Development of commercial supersonic aircraft has been hindered by many related factors including fuel-efficiency, economics, and sonic-boom signatures that have prevented over-land flight. Materials, propulsion, and flight control technologies have developed to the point where, if over-land flight were made possible, a commercial supersonic transport could be economically viable. Computational fluid dynamics, and modern optimization techniques enable designers to reduce the boom signature of candidate aircraft configurations to acceptable levels. However, propulsion systems must be carefully integrated with these low-boom configurations in order that the signatures remain acceptable. One technique to minimize the downward propagation of waves is to mount the propulsion systems above the wing, such that the wing provides shielding from shock waves generated by the inlet and nacelle. This topmounted approach introduces a number of issues with inlet design and performance especially with the highly-swept wing configurations common to low-boom designs. A 1.79%-scale aircraft model was built and tested at the NASA Glenn Research Center's 8-by 6-Foot Supersonic Wind Tunnel (8x6 SWT) to validate the configuration's sonic boom signature. In order to evaluate performance of the top-mounted inlets, the starboard flow-through nacelle on the aerodynamic model was replaced by a 2.3%-scale operational inlet model. This integrated configuration was tested at the 8x6 SWT from Mach 0.25 to 1.8 over a wide range of angles-of-attack and yaw. The inlet was also tested in an isolated configuration over a smaller range of angles-of-attack and yaw. A number of boundary-layer bleed configurations were investigated and found to provide a substantial positive impact on pressure recovery and distortion. Installed inlet performance in terms of mass capture, pressure recovery, and distortion over the Mach number range at the design angle-of-attack of 4-degrees is presented herein and compared to that at 0- degrees, as well as the isolated inlet configuration to highlight installation effects. Performance of the installed inlet fell below that of the isolated inlet at Mach numbers of 1.4 and greater. The installed inlet demonstrated adequate operability over the expected range of angles-of-attack and yaw, but did exhibit definite angle-ofattack and yaw limits at supersonic conditions. At each supersonic flight Mach number, performance parameters near zero yaw angle were relatively insensitive to yaw, but in general the yaw angle yielding best performance was non-zero and varied with angle-of-attack. Performance of the installed inlet is also presented as functions of angle-of-attack and yaw to highlight these effects. Distortion at the aerodynamic interface plane ranged between 10 and 25% at the inlet critical points over the range of flight Mach numbers tested and did not decrease significantly for the isolated inlet. Although these distortion levels would be considered high for operation with a turbine engine, the over-wing installation is likely not as significant a contributor as the low test Reynolds number. This is demonstrated by comparing CFD analysis of the isolated inlet at test scale with that at intermediate and full scales.

Performance↗

NASA NTRS2013

Investigation of Materials for Boundary Layer Control in a Supersonic Wind Tunnel

During operation of the NASA Glenn Research Center 15- by 15-Centimeter Supersonic Wind Tunnel (SWT), a significant, undesirable corner flow separation is created by the three-dimensional interaction of the wall and floor boundary layers in the tunnel corners following an oblique-shock/ boundary-layer interaction. A method to minimize this effect was conceived by connecting the wall and floor boundary layers with a radius of curvature in the corners. The results and observations of a trade study to determine the effectiveness of candidate materials for creating the radius of curvature in the SWT are presented. The experiments in the study focus on the formation of corner fillets of four different radii of curvature, 6.35 mm (0.25 in.), 9.525 mm (0.375 in.), 12.7 mm (0.5 in.), and 15.875 mm (0.625 in.), based on the observed boundary layer thickness of 11.43 mm (0.45 in.). Tests were performed on ten candidate materials to determine shrinkage, surface roughness, cure time, ease of application and removal, adhesion, eccentricity, formability, and repeatability. Of the ten materials, the four materials which exhibited characteristics most promising for effective use were the heavy body and regular type dental impression materials, the basic sculpting epoxy, and the polyurethane sealant. Of these, the particular material which was most effective, the heavy body dental impression material, was tested in the SWT in Mach 2 flow, and was observed to satisfy all requirements for use in creating the corner fillets in the upcoming experiments on shock-wave/boundary-layer interaction.

Braafladt, Alexander↗

NASA NTRS2013

Effects of Hybrid Flow Control on a Normal Shock Boundary-Layer Interaction

Hybrid flow control, a combination of micro-ramps and steady micro-jets, was experimentally investigated in the 15x15 cm Supersonic Wind Tunnel at the NASA Glenn Research Center. A central composite design of experiments method, was used to develop response surfaces for boundary-layer thickness and reversed-flow thickness, with factor variables of inter-ramp spacing, ramp height and chord length, and flow injection ratio. Boundary-layer measurements and wall static pressure data were used to understand flow separation characteristics. A limited number of profiles were measured in the corners of the tunnel to aid in understanding the three-dimensional characteristics of the flowfield.

Hirt, Stefanie M.↗

NASA NTRS2012

Experimental Study of Boundary Layer Flow Control Using an Array of Ramp-Shaped Vortex Generators

The objective of this study was to obtain a database on the flowfield past an array of vortex generators (VGs) in a turbulent boundary layer. All testing was carried out in a low speed wind tunnel with a flow velocity of 29 ft/sec, giving a Reynolds number of 17,500 based on the width of the VG. The flowfield generated by an array of five ramp-shaped vortex generators was examined with hot wire anemometry and smoke flow visualization. The magnitude and extent of the velocity increase near the wall, the penetration of the velocity deficit into the core flow, and the peak streamwise vorticity are examined. Influence of various parameters on the effectiveness of the array is considered on the basis of the ability to pull high momentum fluid into the near wall region.

Hirt, Stefanie M.↗

NASA NTRS2012

Experimental Investigation of Normal Shock Boundary-Layer Interaction with Hybrid Flow Control

Hybrid flow control, a combination of micro-ramps and micro-jets, was experimentally investigated in the 15x15 cm Supersonic Wind Tunnel (SWT) at the NASA Glenn Research Center. Full factorial, a design of experiments (DOE) method, was used to develop a test matrix with variables such as inter-ramp spacing, ramp height and chord length, and micro-jet injection flow ratio. A total of 17 configurations were tested with various parameters to meet the DOE criteria. In addition to boundary-layer measurements, oil flow visualization was used to qualitatively understand shock induced flow separation characteristics. The flow visualization showed the normal shock location, size of the separation, path of the downstream moving counter-rotating vortices, and corner flow effects. The results show that hybrid flow control demonstrates promise in reducing the size of shock boundary-layer interactions and resulting flow separation by means of energizing the boundary layer.

Vyas, Manan A.↗

NASA NTRS2011

Axisymmetric Calculations of a Low-Boom Inlet in a Supersonic Wind Tunnel

This paper describes axisymmetric CFD predictions made of a supersonic low-boom inlet with a facility diffuser, cold pipe, and mass flow plug within wind tunnel walls, and compares the CFD calculations with the experimental data. The inlet was designed for use on a small supersonic aircraft that would cruise at Mach 1.6, with a Mach number over the wing of 1.7. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center in the fall of 2010 to demonstrate the performance and stability of a practical flight design that included a novel bypass duct. The inlet design is discussed here briefly. Prior to the test, CFD calculations were made to predict the performance of the inlet and its associated wind tunnel hardware, and to estimate flow areas needed to throttle the inlet. The calculations were done with the Wind-US CFD code and are described in detail. After the test, comparisons were made between computed and measured shock patterns, total pressure recoveries, and centerline pressures. The results showed that the dual-stream inlet had excellent performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a large stable operating range. Predicted core recovery agreed well with the experiment but predicted bypass recovery and maximum capture ratio were high. Calculations of offdesign performance of the inlet along a flight profile agreed well with measurements and previous calculations.

Chima, Rodrick V.↗

NASA NTRS2011

Experimental Investigation of a Large-Scale Low-Boom Inlet Concept

A large-scale low-boom inlet concept was tested in the NASA Glenn Research Center 8- x 6- foot Supersonic Wind Tunnel. The purpose of this test was to assess inlet performance, stability and operability at various Mach numbers and angles of attack. During this effort, two models were tested: a dual stream inlet designed to mimic potential aircraft flight hardware integrating a high-flow bypass stream; and a single stream inlet designed to study a configuration with a zero-degree external cowl angle and to permit surface visualization of the vortex generator flow on the internal centerbody surface. During the course of the test, the low-boom inlet concept was demonstrated to have high recovery, excellent buzz margin, and high operability. This paper will provide an overview of the setup, show a brief comparison of the dual stream and single stream inlet results, and examine the dual stream inlet characteristics.

Hirt, Stefanie M.↗

NASA NTRS2010

Micro-Ramp Flow Control for Oblique Shock Interactions: Comparisons of Computational and Experimental Data

Computational fluid dynamics was used to study the effectiveness of micro-ramp vortex generators to control oblique shock boundary layer interactions. Simulations were based on experiments previously conducted in the 15 x 15 cm supersonic wind tunnel at NASA Glenn Research Center. Four micro-ramp geometries were tested at Mach 2.0 varying the height, chord length, and spanwise spacing between micro-ramps. The overall flow field was examined. Additionally, key parameters such as boundary-layer displacement thickness, momentum thickness and incompressible shape factor were also examined. The computational results predicted the effects of the micro-ramps well, including the trends for the impact that the devices had on the shock boundary layer interaction. However, computing the shock boundary layer interaction itself proved to be problematic since the calculations predicted more pronounced adverse effects on the boundary layer due to the shock than were seen in the experiment.

Hirt, Stefanie M.↗

NASA NTRS2010

Micro-Ramps for External Compression Low-Boom Inlets

The application of vortex generators for flow control in an external compression, axisymmetric, low-boom concept inlet was investigated using RANS simulations with three-dimensional (3-D), structured, chimera (overset) grids and the WIND-US code. The low-boom inlet design is based on previous scale model 1- by 1-ft wind tunnel tests and features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. Validation of the methodology was first performed for micro-ramps in supersonic flow on a flat plate with and without oblique shocks. For the inlet configuration, simulations with several types of vortex generators were conducted for positions both upstream and downstream of the terminating normal shock. The performance parameters included incompressible axisymmetric shape factor, separation area, inlet pressure recovery, and massflow ratio. The design of experiments (DOE) methodology was used to select device size and location, analyze the resulting data, and determine the optimal choice of device geometry. The optimum upstream configuration was found to substantially reduce the post-shock separation area but did not significantly impact recovery at the aerodynamic interface plane (AIP). Downstream device placement allowed for fuller boundary layer velocity profiles and reduced distortion. This resulted in an improved pressure recovery and massflow ratio at the AIP compared to the baseline solid-wall configuration.

Rybalko, Michael↗

NASA NTRS2010

Progress on a Rayleigh Scattering Mass Flux Measurement Technique

A Rayleigh scattering diagnostic has been developed to provide mass flux measurements in wind tunnel flows. Spectroscopic molecular Rayleigh scattering is an established flow diagnostic tool that has the ability to provide simultaneous density and velocity measurements in gaseous flows. Rayleigh scattered light from a focused 10 Watt continuous-wave laser beam is collected and fiber-optically transmitted to a solid Fabry-Perot etalon for spectral analysis. The circular interference pattern that contains the spectral information that is needed to determine the flow properties is imaged onto a CCD detector. Baseline measurements of density and velocity in the test section of the 15 cm x 15 cm Supersonic Wind Tunnel at NASA Glenn Research Center are presented as well as velocity measurements within a supersonic combustion ramjet engine isolator model installed in the tunnel test section.

Mielke-Fagan, Amy F.↗

NASA NTRS2009

Experimental Investigation of the Application of Microramp Flow Control to an Oblique Shock Interaction

The effectiveness of microramp flow control devices in controlling an oblique shock interaction was tested in the 15- by 15-Centimeter Supersonic Wind Tunnel at NASA Glenn Research Center. Fifteen microramp geometries were tested varying the height, chord length, and spacing between ramps. Measurements of the boundary layer properties downstream of the shock reflection were analyzed using design of experiments methods. Results from main effects, D-optimal, full factorial, and central composite designs were compared. The designs provided consistent results for a single variable optimization.

Hirt, Stefanie M.↗

NASA NTRS2008

CFD Results for an Axisymmetric Isentropic Relaxed Compression Inlet

The OVERFLOW code was used to calculate the flow field for a family of five relaxed compression inlets, which were part of a screening study to determine a configuration most suited to the application of microscale flow control technology as a replacement for bleed. Comparisons are made to experimental data collected for each of the inlets in the 1- by 1-Foot Supersonic Wind Tunnel at the NASA Glenn Research Center (GRC) to help determine the suitability of computational fluid dynamics (CFD) as a tool for future studies of these inlets with flow control devices. Effects on the wind tunnel results of the struts present in a high subsonic flow region accounted for most of the inconsistency between the results. Based on the level of agreement in the present study, it is expected that CFD can be used as a tool to aid in the design of a study of this class of inlets with flow control.

Hirt, Stefanie M.↗