AIAA-93-1924 CFD Applications in an Aeropropulsion Test Environment Greg D. Power and Bonnie D. Heikkinen Sverdrup Technology, Inc., AEDC Group Arnold Air Force Base, Tennessee

AI ANSA EIASM BAS€€ 29th Joint Propulsion

Conference and Exhibit June 28-30, 1993 / Monterey, CA

for permlsslon to copy or republlsh, contact the Amerlcan lnstltute of Aeronautlcs and Astroneutlcs 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

CFD APPLICATIONS IN AN AEROPROPULSION TEST ENVIRONMENT*

Greg D. Power” and Bonnie D. Heikkinen Sverdrup Technology, Inc., AEDC Group

Arnold Engineering Development Center Arnold Air Force Base, Tennessee 37389

Summary

Aeropropulsion testing provides a unique oppor- tunity for interaction between experimental techni- ques and Computational Fluid Dynamics (CFD). While CFD will not supplant testing of propulsion hardware, in recent years the integration of CFD into the test design and analysis process has increased. Several applications are presented which demonstrate the positive impact of computational analysis techniques on large-scale propulsion testing by reducing risks and costs, identifying anomalous behavior, and ex- plaining fluid dynamic phenomena in conjunction with the test data. The test articles range from subscale supersonic nozzles to full-scale commercial high-by- pass engines. Based on the success of these applica- tions, increasing partnerships between experimental and computational fluid dynamics are expected in the future.

v Introduction

The Engine Test Facility (ETF) at the Arnold Engineering Development Center (AEDC) is an Air Force operated national center for ground-base aero- propulsion testing and evaluation. The ETF consists of several large altitude test cells, as well as a multi- tude of small, versatile research cells. Air-breathing engines can be tested with thrust ratings up to 100,000 Ib and at simulated altitudes up to 100,000 ft. Similarly, rocket motors can be tested at design altitudes in either vertical or horizontal test stands. Other test operations include a free-jet facility to simulate enginelinlet interactions and flight environ- ment testing.

Due to the rapid development and improvement of CFD analyses in the past few years coupled with the improved performance of modern computers, the scope of fluid dynamic problems, which can be routinely addressed using numerical techniques, has

increased dramatically. At the same time, propulsion systems have become increasingly complex as im- provements in performance or other mission criteria push the limits of present technologies, requiring de- tailed understanding of component performance and test cell/test article interactions. Thus, Computational Fluid Dynamics is growing into a role which comple- ments and enhances all phases of system design, in- cluding test and evaluation. In fact, CFD is becoming an integral part of the hardware testing process by providing information for pretest facility designlinstru- mentation and posttest analysis,l,* thus reducing risks and improving facility capabilities.

CFD has been developed3.4 and used extensively at AEDC to compliment and enhance the testing capa- bilities in the ETF. In the present paper, a synopsis of current CFD capabilities at the ETF of AEDC is pre- sented followed by several application examples repre- sentative of the diverse and challenging problems currently being addressed.

Analytical Capabilities

In an aeropropulsion test environment, such as that at AEDC, routine analysis of complex flow fields is required. The flow conditions range from low speed, nearly incompressible, coflowing streams to super- sonic jet flows, as well as hypersonic applications. Due to three-dimensional, complex configurations. and generally high Reynolds numbers, non-trivial turbulent flows are common. In addition, chemical kinetics and transport of chemical species are pre- sent, and often dominant, in many propulsion applica- tions. Generally, it is not possible, or practical, for a single CFD analysis to treat the entire range of pro- blems encountered.

There are numerous fluid dynamic analyses being actively pursued at AEDC ranging from quasi- one-dimensional analyses to full Navier-Stokes pro-

“The research reported herein was performed by the Arnold Engineering Development Center (AEDC). Air Force Materiel Command. Work and analysis for this research were done by personnel of Sverdrup Technology, Inc., AEDC Group, technical Sewices contractor for the AEDC propulsion test facilities. Further reproduction is authorized to satisfy needs of the U. S. Government.

“Senior Member, AIAA. ”’ This papensdeclareda workoftheU.S. government and i s notsubjedtocopytigh? protection i n the United Stater.

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grams with chemistry. The quasi-one-dimensional Dynamic Turbine Engine Compressor Code (DYNTECC),S developed in conjunction with Wright Laboratories Compressor Research Facility, provides a reliable approach to propulsion component analysis, complementing the less empirical Navier-Stokes (NS) codes. The PARC code,$ the primary CFD tool in use for engine test analysis at AEDC, is a mature, well- established full Navier-Stokes code which has evolved to include numerous general boundary conditions, grid blocking, and robust solution techniques. As a conse- quence of a recent alliance between the NASA Lewis Research Center and the AEDC. PARC is evolving to reflect the widespread use of the code in government and industry with a name change to reflect this evolu- tion (NPARC - National PARC). The GASP codes in- cludes recent developments such as an upwind and a space-marching algorithm. For nozzle flow fields in which species transport and chemical kinetics are important and the viscous flow is restricted to a thin layer, the TDWBLM code7.s is used extensively. For more general chemically reacting flows, several analysis codes are being evaluated including TUFF,g KIVA,lo RPLUS,11 GIFS12 and GASP. Considerable code development is also on-going to continually im- prove the CFD programs to meet the changing applica- tion needs.

The fluid dynamic analysis programs provide the backbone of any CFD effort. However, the supporting pre- and post-processing programs have grown in im- portance as problem complexity has increased. Pre- processing includes geometry definition, grid genera- tion, and the generation of flow solver input. INGRID, an interactive grid generation program derived from SVTGD,13 is being replaced by GRIDGEN.14 with which even the most complex geometrical problems can be modeled efficiently with significantly less user effort. In addition to grid generation, GRIDGEN V6.0 has been modified at AEDC to provide boundary condi- tion information for the PARC code, thus reducing the time consuming task of constructing complex input files. To compliment the PARC input interface in GRIDGEN, a graphical user interface has been deve- loped to aid the user in constructing and editing the PARC input files.

In addition to the determination of performance characteristics of rocket and turbine engines, the alti- tude test cells at AEDC provide an opportunity to ac- quire hot-part and plume radiation data over a range of operating conditions. To support these efforts, analytical tools such as the Standardized Infrared Radiation Model (SIRRM), developed as the Joint- Army-Navy-NASA-Air Force (JANNAF) standard, are being utilized extensively. An upgraded model, Infrared Radiation Model User Friendly (IRRMUF),

was developed at AEDC. A description of the diag- nostic and analytical techniques can be found in Ref. 15.

L/ Applications

In this section, several applications related to aero- propulsion testing at AEDC are presented. For each application a discussion of the test is given along with a discussion of the particular issue to be addressed using CFD. A synopsis of the CFD model and solution procedure is presented next followed by a discussion of the significant CFD results and any comparison with data. Finally, a brief discussion of the impact of the CFD analysis on the resolution of the original issue is given.

Free-jet Nozzle Flow Quality

The C2 subsonic free-jet facility in the ETF at AEDC provides the capability to study full-scale articles at realistic Reynolds numbers and Mach numbers. A schematic of the C2 installation is shown in Fig. 1 . The study of the interaction between the full-scale engine, inlet, and forebody of an aircraft is one appli- cation of this facility, as described in a following section. Throughout the testing of the C2 subsonic free-jet nozzle, unexplained variations in nozzle exit Mach number and flow angle persisted over a range of Mach numbers, nozzle pitch angles, and yaw angles. A com- putational nozzle flow quality study was initiated to determine if these variations could be reproduced in an attempt to determine their origin.

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The flow field was simulated using the PARC code, where the effects of turbulence were modeled using a low Reynolds number form of the (k- E) turbulence model.4 The initial computational effort focused on 2D symmetry plane calculations in order to study the detailed flow structure. A 2D grid consisting of appro- ximately 10,000 nodes was used to discretize the geometry and flow field for the symmetry plane simulations. The 3D configuration was also simulated to assess the impact of 3D flow features on the 2D simulations and to identify any prominent 3D flow structures. Approximately 500,000 nodes were used to discretize the 3D flow field. For the present study, a nozzle exit Mach number of 0.8 was simulated at an inlet total pressure of 3.26 psia and total tem- perature of 461.7"R.

Figure 2 shows the results of the 2D simulations on the symmetry planes for the "as built" configura- tion. The Mach number contours in the XY plane indicate that the flow separates as it accelerates around the sharp nozzle leading edge and re-attaches at approximately the point where the nozzle becomes constant area. This extended development length for

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the free shear layer, created by the separation and reattachment in a region where minimal mean flow acceleration is available to reaccelerate this flow, allows the development of a significant region 01 degraded total pressure. This gradient in total pres- sure and hence Mach number persists downstream resulting in a significant gradient in Mach number at the nozzle exit plane. For the XZ symmetry plane, while inlet lip separation is present, it is not as severe, with reattachment occurring well upstream of the point where the nozzle becomes constant area. Such a reattachment location provides an opportunity for reacceleration of the reattached flow, minimizing any near wall total pressure loss and hence Mach number gradients.

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The above results indicate that inlet lip separation represents a likely cause for the more significant nozzle exit plane nonuniformities. To further study this effect, modifications of the inlet lip radius were investigated. The original nozzle lip geometry had a lip entrance radius of 0.38 in. to simulate the actual nozzle wall thickness. The lip radius was enlarged to 7.8 in. in the XY plane, since the effects of separa- tion were most pronounced in this plane. This con- figuration change eliminated flow separation at the nozzle inlet lip and considerably reduced the flow nonuniformities at the nozzle exit, as shown by the Mach number contours in Fig. 3.

W The effect of the inlet lip modification on the

nozzle exit flow field can be seen in the Mach num- ber profiles at the exit plane shown in Fig. 4 for both the original and modified inlet lip geometries. Based on the 2D calculations, the shear layer in the modi- lied geometry is confined to the near-wall region, resulting in a more uniform Mach Number profile. A 50% reduction in lip radius (3.9 in.) again resulted in significant flow separation, as in the original geo- metry. Data obtained at a 15-deg nozzle pitch verifies the trend predicted by the 2D simulation, but the absolute distribution is not well presented.

To compliment the 2D solutions discussed above, a 3D 114 plane of symmetry solution was computed to investigate the spanwise character of the inlet lip separations and their effect on the nozzle exit plane nonuniformities. Based on the geometry, the assump- tion of a vertical plane of symmetry is valid. The assumption of a horizontal plane of symmetry, while not strictly consistent with the geometry, is con- sidered justified by the considerable vertical span of the nozzle relative to the wall boundary layers. The predicted exit Mach Number profiles shown in Fig. 4 compare quite well with the measured distribution at

,-, 15-deg pitch. These results indicate that the 2D simulations are conservative for design purposes.

Representative results of the 3D simulation are shown in Fig. 5. An isosurface of a constant value of negative streamwise velocity indicates large 'bubbles' representing inlet separation similar to that predicted in the 2D simulations. Consistent with the 2D results, there is separation in both the symmetry planes, but it is more pronounced on the XY symmetry plane. In addition, the separations disappear completely as the interior corner of the nozzle is approached.

Based on the CFD simulations, a subscale test is underway in which the nozzle inlet lip radius was increased to an effective 10.5-in. radius to be certain of eliminating the inlet separation.

Icing Spray Bar Analysis

Ice accumulation on propulsion system surfaces generally results in a degradation of performance and operational safety. The use of altitude test cells has become an acceptable approach for evaluation of the effects of icing conditions on these systems.16 At AEDC an existing facility is being modified to include icing spray bars upstream of the engine test section. There are several considerations in designing the spray bars, including the uniformity of the icing pattern and the flow blockage effects of the icing bars. Based on the consideration of icing uniformity, the original de- sign stipulated 17 spray bars placed horizontally across the existing test cell duct. To determine the effect of this design on cell blockage, a 2D CFD simulation was performed.

Taking advantage of symmetry, one-half of a spray passage was modeled with symmetry boundary conditions along the centerline of the spray bar and along the center of the gap between spray bars. The PARC code was run inviscid to simulate the worst case scenario, i.e., no boundary layer blockage. The flow conditions were specified for an simulated altitude of 20,000 ft and flight Mach number of 0.48. These flow conditions were chosen to correspond to a case in which choking was most likely.

Figure 6 shows that the original spray bar grid was placed within the existing constant area section up- stream of the test section. The computed Mach num- ber contours for this case are shown in Fig. 7 for a single spray bar. The simulation for this configuration at the specified conditions does indicate choked flow within the spray passage. To remedy this situation, the test cell was redesigned to place the spray bar grid within a converging/diverging section, as shown in Fig. 6, such that the gap between spray bars could be increased while maintaining the same number of bars.

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Forebody Design

The interaction between an aircraft propulsion system, the engine inlet, and the airframe must be properly accounted for in any propulsion system design. This interaction is particularly important for fighter aircraft in which the engines are closely coup- led to the airframe, and large angles of pitch and/or yaw are common. The free-jet facility in the AEDC Aero- propulsion Systems Test Facility (ASTF) is designed to test near full-scale engine inlets in the presence of the aircraft forebody. Often the forebody must be placed within the free-jet nozzle, requiring a forebody simulator (FBS) that will fit inside the nozzle while accurately siinulating the flow at the engine interface (Fig. 8). Trial and error design of the FBS would be quite expensive and time consuming, requiring several FBS configurations and multiple tests. It was recogni- zed that the design of the test hardware and test con- ditions to accurately simulate the forebody interaction could be more efficiently accomplished through CFD simulations.

As part of a multiphase experimental program, a 116-scale F-15 inlet model was tested in a pilot sub- scale free-jet facility with four different forebody simulators. Tests were conducted at subsonic Mach numbers of 0.3, 0.6, and 0.9 for angles-of-attack from -10 to 35 deg and sideslip angles of -10 to 35 deg and at a supersonic Mach number of 2.2 for angles-of-attack from -5 to -15 deg and sideslip angles from -6 to 6 deg.

As a step toward applying CFD to design fore- body simulators, the PARC code is being used to simu- late free-jet test points with the objective being a calibrated code for free-jet applications.1' To verify the effectiveness of the FBS, the flow conditions at a predetermined inlet reference plane (IRP),either measured during the test or determined using CFD, are compared to those experienced in free flight, in this case subscale wind tunnel data. CFD would be used to vary the FBS and the test conditions until the IRP conditions are reproduced.

The most recent calibration of the PARC code has beenforaset of test points in the supersonic flight re- gime at a Mach number of 2.2.17 Figure 9 shows the computational model of the test hardware. The forebody simulator was derived by taking a slice of the actual airframe fuselage. A Chimera grid scheme,l8.19 in which the FBS grid is embedded within the free-jet facility grid, was used to minimize the grid development time required to properly set up the model for different test conditions.

The free-jet test conditions were imposed by speci- fying test cell total pressure and total temperature at the inflow plane of the nozzle, and static cell pres- sure at the downstream exit plane. The nozzle exit Mach number was set by varying the throat area analogous to the experimental hardware. The inlet flow was specified by imposing a mass flux boundary condition at the downstream exit plane of the inlet. All solid walls were modeled as "slip" boundaries (Le., no boundary-layer calculation). The solution algorithm was run viscous/turbulent, using the Thomas algebraic model3 to capture the free-jet shear layer. Approximately half a million grid points were distri- buted over 14 zonal blocks. Because of the arbitrary boundary shapes generated by the Chimera grid scheme, an explicit solution algorithm was required which results in slower convergence compared to the nowchimera version of the PARC code.

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Three supersonic Mach 2.2 computations were performed, two at high angle-of-attack (12 deg) with sideslip angles of 3 and 0 deg and one at zero angle- of-attack. Figures 10 and 11 show the computational results and test comparison for the center IRP rake position, shown as a solid line in Fig. 10. The com- plex nature of the flow field is evident from the Mach number contour plot in Fig. 10 as the bow shock from the forebody simulator reflects off of the lower nozzle wall and intersects the oblique shock from the upper inlet cowl lip. The Mach number along the IRP shown in Fig. 11 is compared to measured subscale data indicating that CFD can predict the fore- bodylinleffengine interactions. This capability in con- junction with recent developments in inverse design technology20 could lead to decreased design time for forebody simulators.

Subscale Supersonic Nozzle

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For most experimental studies, it is necessary to determine not only what quantity should be mea- sured, but also where to place the instrumentation to capture the most interesting features or to acquire the most pertinent information. Thus, an a priori knowledge of the flow field is often required. For the study of the shock structure within supersonic nozzle flows, the location of the Mach disks and plume dia- meter is necessary to properly place the instrumenta- tion. In a recent study by AEDC personnel at NASA Stennis.21 the flow within a subscale supersonic nozzle was studied experimentally, and CFD was applied to provide guidance for the placement of instrumentation. The geometry and flow field was discretized using approximately 50,000 nodes. The chamber pressure of the nozzle was 500 psiaat a tem- perature of 6,OOO'R for a design Mach number of 'L,'

2.9. The internal nozzle flow was modeled up to the ,hroat using the TDK code. Assuming frozen chemistry downstream of the throat, the nozzle and jet flows were simulated, starting from the TDK solution with a constant ratio of specific heats of 1.2. The axisymmetric PARC code was applied using the Thomas algebraic turbulence model.

The predicted Mach number contours are shown in Fig. 12 compared to a photograph taken while the motor was firing. As indicated in this figure, the CFD results properly predicted the location of the Mach disks and provided the necessary information for proper instrumentation placement. The location of the first Mach disk was predicted to be 2.29 in. from the nozzle exit, within 0.08 in. of the measured location.

Scramjet Probe Design

A detailed understanding of the flow field within a scramjet engine is an important requirement in the successful design of the NASP. Due to the hostile en- vironment within a scramjet, probes must be carefully designed to withstand the extreme temperatures and pressures. Due to the recent successful experience at AEDC with water-cooled Mach flow angularity (MFA) probes, asmallerversion was requested for use in the NASP Concept Demonstrator Engine (CDE) planned for the 8-ft High Temperature Tunnel (HTT) at NASA Langley Research Center. Figure 15 shows a photo- graph of the MFA probe. Although the design was based on MFA probes currently in use, questions arose during the design validation testing that could not be resolved empirically. The two major concerns were related to the location of .the cone static pressure orifices and the impact of the rounding of sharp corners due to erosion. An axisymmetric simula- tion was performed using the PARC code. The flow field was simulated over a matrix of flow conditions ranging from Mach numbers of 1.75 to 6.0 over a range of conditions corresponding to expected loca- tions within the engine for both the sharp and rounded corner configurations. To calibrate the PARC code simulations, the results for two cases were compared to experimental data obtained at the VKFIACL facility at AEDC.22 The two cases correspond to local Mach numbers of 1.788 and 4.0. The results of the cali- bration cases are shown in Fig. 14, a plot of probe surface static pressure normalized by the total pressure behind the normal shock. These results indicate excellent agreement with the experimental data at the static orifice location.

The matrix of test cases confirmed that the loca- tion of the static pressure orifice is sufficient for all conditions studied in that the orifice was not affected by the bow shock at the probe tip or the shoulder expansion. In addition, the effect of rounding the

shoulder was found to be negligible. The CFD results provided confidence in the integrity of the original probe design. More details of the probe design and CFD results can be found in Ref. 22.

Hydrogen Disposition

As the design of NASP-type vehicles progresses, the use of existing facilities to test the proposed scram- jet propulsion systems is being investigated. A major concern in the use of ground-based facilities is the possibility that testing of these hydrogen-fueled pro- pulsion systems can result in the accumulation of ex- plosive exhaust gases during transient off-design operation or during emergency shut down pro- cedures. One proposed method to eliminate this explosive hazard is to burn off the combustible ex- haust mixture. A proposed system to burn the exhaust gas consists of a series of "V-gutters" placed per- pendicular to the gas flow across the exhaust duct. Equipped with a number of hydrogen torches, this stabilized bluff body combustion system is expected to eliminate excess hydrogen. Because of the size of the test cell required for the propulsion testing and associated hazards, it was decided to validate the exhaust burning system design using a stream tube model operating in a research test cell. Numerical modeling of the exhaust disposition system is de- sirable to understand the physics of the streamtube system and to extend the experimental results to the actual large cross section propulsion test cell.

Validation of the numerical model was desired prior to obtaining the streamtube combustion data. Unfortunately, experimental data for conditions appropriate to the test conditions were not found for hydrogen burning. A detailed data set was, however, found for stabilized premixed propane combustion behind a rearward facing step, so it was decided to compare the numerical model with this data.23 The numerical model chosen to simulate stabilized com- bustion was the KIVA-II computer code developed at LOS Alamos Scientific Laboratories.10 Code modifica- tion was required to place the rearward facing step in the flow, to set the boundary conditions on the step, and to ignite the gaseous mixture. The quasi-global hydrocarbon oxidation reaction set reported by West- brook and Dryer24 was employed for this computa- tion with the backward reaction rates determined by the reaction equilibrium constants.

Data obtained for this experiment included mea- surements of velocity, turbulent intensity, and tem- perature for both reacting and non-reacting flow. Species concentration measurements were also taken for the reacting flow. Cross-stream measurements were obtained at a number of streamwise locations to provide profile plots of the data. Photographs of the

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turbulent structure were also taken to indicate eddy size and frequency. Comparisons of computed mean velocity and turbulent intensity with data for cold flow were very encouraging. Hot-flow comparisons are in progress, and a contour plot of the computed tem- perature field is shown in Fig. 15. The undulations in the contours indicate that the solution is unsteady due to vortex shedding off the step. Comparison with the schlierin photograph of the turbulent flow field indicates qualitative agreement with the experimental data. Based on the confidence in the solution pro- cedure provided by these results, simulation of the streamtube model is currently being pursued to provide pretest information on the effectiveness of the hydrogen disposition system.

Compressor Analysis

In addition to overall engine performance, com- ponent performance is also an important aspect of testing that can benefit from reliable analytical techni- ques.The Compressor Research Facility at Wright- Patterson Air Force Base will be testing the ADLARF (Augmented Damping of Low Aspect Ratio Fans) fan, and one of the objectives of the test program will be to evaluate five different casing treatments over the first-stage rotor. From a previous test of the ADLARF fan, it was determined that the second stage was the critical stalling stage over a wide range of speeds. In arecent investigation the DYNTECC code, a dynamic compression system model, has been calibrated for the ADLARF two-stage, low-aspect ratio fan.25 After calibration, the program was used to perform a para- metric study to determine how casing treatments effect the stage and overall stalling airflow capability of the fan. This study focused on showing what im- provements were necessary to the second stage stalling airflow to realize gains from applying various casing treatment to the first stage.

DYNTECC uses stage characteristics based on aerodynamic performance data. To model the effect of casing treatments, data from several investigations reported in the literature were used to calibrate the DYNTECC model. The calibrated model showed that applying casing treatment in combination with changing the variable vanes is an effective way to lower the stage stalling airflow. OYNTECC predicted, as shown in Fig. 16, that these methods will need to provide as much as a 14-percent improvement in the second-stage stalling airflow capability to get a 20- percent increase in stalling airflow capability from the first stage. In addition, the model showed that casing treatments are an effective means of overcoming stall airflow losses due to distortion. More details of the analysis and results can be found in Ref. 25.

Aeroacoustlc Analysis Support

Turbineenginesare tested in much the same way as rocket engines in that the external pressure is simu- lated by the combined effects of facility compressors and diffuser pumping. Due to the close coupling of the facility and the test article, undesirable acoustic or fluid dynamic interactions can occur. Noise can also be a factor during altitude testing when the reliability of the measurements or the safety of the test article or the facility are jeopardized due to an anomalous acoustic environment. An example of such a case is a recent full-scale test conducted in the C2 test facility at AEDC in which the supersonic exhaust of a turbine engine produced high-intensity discrete acoustic tones.26 A small-scale experimental test pro- gram and acoustic analysis were performed at the Georgia Tech Research Institute (GTRI) by Ahuja et. alz7 to investigate the source of the noise.

The experimental apparatus was a concentric axisymmetric nozzle/diffuser combination enclosed by a test cell with the facility exhaust ducting simulated for approximately four diameters (Fig. 17). One of the principal concerns relative to the fluid dynamic per- formance of this configuration was whether the nozzle plume attached to the facility exhaust duct wall down- stream of the diffuser prior to the end of the exhaust duct. The motivation for this concern was driven by the importance of the recirculating flow field in the diffuser. If the plume does not attach, a significant amount of flow may be drawn in at the exit of the exhaust duct to satisfy the entrainment requirements of the jet. Should this situation exist, the mean tem- perature of the flow in the diffuser surrounding the jet could be substantially different from that which would exist if attachment had taken place. This in turn would significantly alter the acoustic resonant res- ponse of the diffuser with a resulting impact on the acoustic behavior of the overall system.

As a means of investigating this phenomenon, an axisymmetric solution was acquired for a nozzle/dif- fuserltest cell configuration closely related to that shown in Fig. 17 but with the exhaust duct extended beyond the point where attachment was expected. This allowed determination of a predicted attachment length for comparison against the known length of the exhaust duct. Figure 18 shows the result of the computed streamlines. The red streamlines represent flow from the nozzle and the blue represent secon- dary cooling air (10 percent of the nozzle flow rate) introduced into the test cell. As indicated, the attach- ment point is at approximately four exhaust duct dia- meters downstream of the nozzle exit which implies that the subscale model should be adequate with respect to jet reattachment.

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Conclusions

An overview of the CFD effort at the Engine Test V Facility at the Arnold Engineering Development

Center indicates the diversity of applications and analyses currently being pursued at AEDC. The appli- cations discussed range from facility design considera- tions to instrumentation design and data analysis. The analytical/numerical approaches range from a quasi-one-dimensional compression system model to full 3D Navier-Stokes codes with chemistry. The re- sults demonstrate the ability of CFD to positively impact aeropropulsion testing operations. The quality and success of these applications indicate that CFD is becoming more accepted as a method to reduce risks and costs associated with engine testing.

Acknowledgments

The CFD applications reported in this paper were performed by Kyle Cooper, Milt Davis, Alan Hale, Bonnie Heikkinen, Ralph Jones, Dennis Lankford, Mike McClure, and Greg Power. The Air Force Pro- ject Manager was Lt. Scott Tennent, AEDCIDOTF.

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20. Huddleston, D. H. and Mastin, C.W. "Optimi- zation of Aerodynamic Designs Using Computational Fluid Dynamics." AGARD Specialists Meeting on Computational Methods for Aerodynamic Design (inverse) and Optimization, Loen, Norway, May 1989.

21. Dawson, M. C., Douglas, F. 111, and Orlin, P. A. "ASRM Plume Deflector Analysis Program." AlAA 17th Aerospace Ground Testing Conference, Nash- ville, TN, July 6-8, 1992.

22. Hiers, R.S. Jr. and Jalbert, P. A. "Mach Flow Angularity Probes for Scramjet Engine Flowpath Diagnostics." Sept, 27-30, 1993 (to be presented at SAE AEROTECH '93 Costa Mesa, CA).

23. Ganji, A.T. and Sawyer, R.F., "Turbulence, Combustion, Pollutant and Stability Characterization of a Premixed, Step Combustor." NASA CR 3230, January, 1980.

24. Westbrook, C.K. and Dryer, F.L. "Simplified Reaction Mechanisms for the Oxidation fo Hydro- carbon Fuels in Flames." Combustion Science and Technology, 1981, Vol. 27, pp.34-43. ..-..

25. Gorrell, S.E. and Davis, M.W., Jr. "Applica- tion of a Dynamic Compression System Model to a Low Aspect Ratio Fan: CASING Treatment and Dis- tortion." A I M 93-1871, AIAWSAWASMWASEE 29th Joint Propulsion Conference, Monterey, CA, June 28- 30, 1993.

26. Jones, R.R. 111, and Lazalier. G.R. "The Acoustic Response of Altitude Test Facility Exhaust Systems to Axisymmetric and Two-Dimensional Ex- haust Plumes." DGLWAIAA 92-02-131, DGLWAIAA 14th Aeroacoustics Conference, May 1992.

27. Ahuja, K.K., Massey, K.C., Flemming, A.J., Tam, C.K.W., and Jones, R.R. Ill, "Acoustic Interac- tions Between an Altitude Test Facility and Jet Engine Plumes - Theory and Experiments." AEDC- TR-91-20 (AD-A245463), January 1992.

Fig. 1, C2 subsonic freejet installation.

8

XZ Fig. 2. Mach contours on symmetry planes - original freejet nozzle geometry.

Fig. 3. Mach contours on XY symmetry plane - 7.8 in. lip radius

9

- ~ . - 20 ORIGINAL GEOMETRY -- 20 MODIFIEO GEOMETRY A I IS OEG PITCH -- 30 ORIGINAL GEOMETRY

MACH NUMBER .~ Fig. 4. Comparison of Mach number profiels at nozzle

exit plane.

Fig. 5. Extent of separation from 3D freejet nozzle simulation

10

-ICING EARS

Fig. 6. Schematic of icing

A - b

spray bar configurations.

Fig. 7. Predicted Mach contours between icing spray bars.

Fig. 8. Forebody simulator installation in freejet nozzle, .. .. ....

Fig. 9. Computational model of forebody simulator installation.

.. ....

Fig. I O . Mach contours at cenerline of engine inlet.

12

f

19’2 16.0 c t

MACH NUMBER Fig. 11. Comparison of predicted and measured

Mach numbers along IRP.

Fig. 12. Predicted Mach contours compared with photograph of DTF motor firing. ._- 13

.. .

-----c--------

CALIBRATION D A f A

- PARC SOLUTIONS

I I 1 0.20 0.40 0.60

Fig. 14. Predicted probe surface pressure distribution X, IN.

Fig. 13. Photograph of MFA probes. compared with data.

Fig. 15. Comparison of predicted temperature contours with Schlieren for premixed propane cobustion.

14

SIABlL l IY LIMIT /- WITHOUT CASING STABILITY LIMIT WITH STAGE I AND 2 CASING rrd

p 4 . h

Y

Y

&

o ORIGINAL B A W N t a MOOIIIfD I B 2 llAGC 1201131

"

4 . 2 120 130 I40 I 5 0 160

COMPRESSOR [ORRECIED AIRFLOW, lbmirer Fig. 16. Dyntecc prediction of tip casing treating

effects on overall compression system performanceioperability.

A X I S Y M M E I R K I- 51.0 -* CONVERGENT NOZZLE 7 rn

Fig. 17. Subscale axisymmetric jevdiffuser configuration.

i

... . . . . ~~. . . . . ~. . .

Fig. 18. Predicted streamline pattern for fullscale jevdiffuser model

....I....

15