ALBERT

Description: The work to be described was performed at the NASA Langley UPWT (4-ft supersonic), test section #2, during 21-24 May 1996. The configuration being tested was the 1.675% Ref H controls model; test conditions were Ma = 2.40, Re = 3 million/ft. This was an exploration of a new technique, and it was not intended to provide definitive comparison of measured and computed skin friction results. It is, however, hoped that the experience gained will make such a rigorous comparison possible in the future.

Description: To summarize the significant highlights in this report: (1) Data quality, determined by multiple repeat runs performed on the TCA baseline configuration, and long-term repeatability, determined by comparing baseline Reference H data from this test to a previous test, have been shown to be good. (2) The longitudinal stability of the TCA is more non-linear than for the Reference H, and while it is similar at normal lift values, the TCA has considerably more pitch-up at higher lift. (3) Longitudinal control effectiveness of the TCA is similar to the Reference H and the ratio of elevator effectiveness to horizontal tail effectiveness is approximately 0.3. 4) The directional stability of the TCA is improved relative to Reference H at higher angles-of attack. The chine is effective for improving directional stability. (5) The directional control effectiveness 'of the TCA rudder is the same as that of the Reference H rudder at low angles-of-attack, after taking factors, such as number of rudder panels deflected and vertical tail volume into account. However, rudder effectiveness was shown to be reduced at higher angles-of-attack. (6) The lateral stability was shown to be reduced relative to the Reference H, which may be beneficial at low speeds for alleviating lateral control saturation. (7) Lateral control effectiveness for the TCA was shown to be similar to the Reference H for negative trailing-edge flap deflections and was reduced by approximately 25% for positive trailing-edge flap deflections.

Description: This paper gives the results of a grid study, a turbulence model study, and a Reynolds number effect study for transonic flows over a high-speed aircraft using the thin-layer, upwind, Navier-Stokes CFL3D code. The four turbulence models evaluated are the algebraic Baldwin-Lomax model with the Degani-Schiff modifications, the one-equation Baldwin-Barth model, the one-equation Spalart-Allmaras model, and Menter's two-equation Shear-Stress-Transport (SST) model. The flow conditions, which correspond to tests performed in the NASA Langley National Transonic Facility (NTF), are a Mach number of 0.90 and a Reynolds number of 30 million based on chord for a range of angle-of-attacks (1 degree to 10 degrees). For the Reynolds number effect study, Reynolds numbers of 10 and 80 million based on chord were also evaluated. Computed forces and surface pressures compare reasonably well with the experimental data for all four of the turbulence models. The Baldwin-Lomax model with the Degani-Schiff modifications and the one-equation Baldwin-Barth model show the best agreement with experiment overall. The Reynolds number effects are evaluated using the Baldwin-Lomax with the Degani-Schiff modifications and the Baldwin-Barth turbulence models. Five angles-of-attack were evaluated for the Reynolds number effect study at three different Reynolds numbers. More work is needed to determine the ability of CFL3D to accurately predict Reynolds number effects.

Description: The NASA High Speed Research (HSR) Program is intended to establish a technology base enabling industry development of an economically viable and environmentally acceptable second generation high speed civil transport (HSCT). The HSR program consists of work directed towards several broad technology areas, one of which is aerodynamic performance. The objective of the Configuration Aerodynamics task of the Aerodynamic Performance technology area is the development of aerodynamic drag reduction, stability and control, and propulsion airframe integration technologies required to support the HSCT development process. Towards this goal, computational and empirical based aerodynamic design tools are being developed, evaluated, and validated through ground based experimental testing. In addition, methods for ground to flight scaling are being developed and refined. Successful development of validated design and scaling methodologies will result in improved economy of operation for an HSCT and reduce uncertainty in full-scale flight predictions throughout the development process.

Description: It is not unusual when comparing CFD data to experimental data to find discrepancies between the results. Sometimes forces and moments compare well, while surface pressures do not, and vice versa. It is commonplace for the researcher to believe that the flow field has been accurately simulated when these types of measurements compare well. However, being able to routinely predict boundary layer transition and separated flows are not guaranteed. In fact accurate simulation of these types of flow physics has been a challenge to the CFD community. In order to improve Navier-Stokes predictions for complex vortical flow fields, more detailed information about the flow physics is necessary. Unfortunately, the many wind-tunnel tests performed in Langley's NTF and 14x22 facilities as well as in the Ames' 12 ft. Tunnel provided little information about the detailed flow physics, and no priority was given to obtaining any CFD measurements. Using the latest experimental techniques, this information can and should be obtained for present and future use.

Description: This paper focuses on the parallel computation of aerodynamic derivatives via automatic differentiation of the Euler/Navier-Stokes solver CFL3D. The comparison with derivatives obtained by finite differences is presented and the scaling of the time required to obtain the derivatives relative to the number of processors employed for the computation is shown. Finally, the derivative computations are coupled with an optimizer and surface/volume grid deformation tools to perform an optimization to reduce the drag of a three-dimensional wing.

Description: Current parallel computational approaches involve distributed and shared memory paradigms. In the distributed memory paradigm, each processor has its own independent memory. Message passing typically uses a function library such as MPI or PVM. In the shared memory paradigm, such as that used on the SGI Origin 2000 machine, compiler directives are used to instruct the compiler to schedule multiple threads to perform calculations. In this paradigm, it must be assured that processors (threads) do not simultaneously access regions of memory in such away that errors would occur. This paper utilizes the latest version of the SGI MPI function library to combine the two parallelization paradigms to perform aerodynamic shape optimization of a generic wing/body.

Description: The primary objectives of this study were to expand the data base showing the effects of LE radius distribution and corresponding sensitivity to Rn at subsonic and transonic conditions, and to assess the predictive capability of CFD for these effects. Several key elements led to the initiation of this project: 1) the necessity of meeting multipoint design requirements to enable a viable HSCT, 2) the demonstration that blunt supersonic leading-edges can be associated with performance gain at supersonic speeds , and 3) limited data. A test of a modified Reference H model with the TCA planform and 2 LE radius distributions was performed in the NTF, in addition to Navier-Stokes analysis for an additional 3 LE radius distributions. Results indicate that there is a tremendous potential to improve high-lift performance through the use of a blunt LE across the span given an integrated, fully optimized design, and that low Rn data alone is probably not sufficient to demonstrate the benefit.

Description: This paper presents results of a study which attempted to provide some understanding of the relationship between skin friction drag estimates produced by flat plate methods and those produced by Navier-Stokes computations. A brief introduction is followed by analysis, including a flat plate grid study, analysis of the wing flow, an analysis of the fuselage flow. Other results of interest are then presented, including turbulence model sensitivities, and brief analysis of other configurations.

Description: Efforts towards understanding boundary layer transition characteristics on a High Speed Civil Transport (HSCT)-class configuration in the National Transonic Facility (NTF) are ongoing. The majority of the High Speed Research (HSR) data base in the NTF has free transition on the wing, even at low Reynolds numbers (Rn) attainable in conventional facilities. Limited data has been obtained and is described herein showing the effects of a conventional, Braslow method based wing boundary-layer trip on drag. Comparisons are made using force data polars and surface flow visualization at selected angles-of-attack and Mach number. Minimum drag data obtained in this study suggest that boundary layer transition occurred very near the wing leading edge by a chord Rn of 30 million. Sublimating chemicals were used in the air mode of operation only at low Rn and low angles-of-attack with no flap deflections; sublimation results suggest that the forebody and outboard wing panel are the only regions with significant laminar flow. The process and issues related to the sublimating chemical technique as applied in the NTF are discussed. Beyond the existing experience, status of efforts to develop a production transition detection system applicable to both air and cryogenic nitrogen environments is presented.

Description: Model deformation measurement techniques have been investigated and developed at NASA's Langley Research Center. The current technique is based upon a single video camera photogrammetric determination of two dimensional coordinates of wing targets with a fixed (and known) third dimensional coordinate, namely the spanwise location. Variations of this technique have been used to measure wing twist and bending at a few selected spanwise locations near the wing tip on HSR models at the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel. Automated measurements have been made at both the Transonic Dynamics Tunnel and at Unitary Plan Wind Tunnel during the past year. Automated measurements were made for the first time at the NTF during the recently completed HSR Reference H Test 78 in early 1996. A major problem in automation for the NTF has been the need for high contrast targets which do not exceed the stringent surface finish requirements. The advantages and limitations (including targeting) of the technique as well as the rationale for selection of this particular technique are discussed. Wing twist examples from the HSR Reference H model are presented to illustrate the run-to-run and test-to-test repeatability of the technique in air mode at the NTF. Examples of wing twist in cryogenic nitrogen mode at the NTF are also presented.

Description: To develop full scale flight performance predictions an understanding of Reynolds number effects on HSCT-class configurations is essential. A wind tunnel database utilizing a 2.2% scale Reference H model in NASA Langley Research Centers National Transonic Facility is being developed to assess these Reynolds number effects. In developing this database temperature and aeroelastic corrections to the wind tunnel data have been identified and are being analyzed. Once final corrections have been developed and applied, then pure Reynolds number effects can be determined. In addition, final corrections will yield the data required for CFD validation at q = 0. Presented in this report are the results of seven tests involving the wing/body configuration. This includes summaries of data acquired in these tests, uncorrected Reynolds number effects, and temperature and aeroelastic corrections. The data presented herein illustrates the successes achieved to date as well as the challenges that will be faced in obtaining full scale flight performance predictions.

Description: Experience with afterbody closure effects and accompanying test techniques issues on a High Speed Civil Transport (HSCT)-class configuration is described. An experimental data base has been developed which includes force, moment, and surface pressure data for the High Speed Research (HSR) Reference H configuration with a closed afterbody at subsonic and transonic speeds, and with a cylindrical afterbody at transonic and supersonic speeds. A supporting computational study has been performed using the USM3D unstructured Euler solver for the purposes of computational fluid dynamics (CFD) method assessment and model support system interference assessment with a focus on lower blade mount effects on longitudinal data at transonic speeds. Test technique issues related to a lower blade sting mount strategy are described based on experience in the National Transonic Facility (NTF). The assessment and application of the USM3D code to the afterbody/sting interference problem is discussed. Finally, status and plans to address critical test technique issues and for continuation of the computational study are presented.

Description: The Boeing Reference H configuration was tested in the NASA Ames 9x7 Supersonic Wind Tunnel. A simulated unstarted inlet was evaluated as well as the aerodynamic performance of the configuration with and without nacelle and diverter components. These experimental results were compared with computational results from the unstructured grid Euler flow solver AIRPLANE. The comparisons between computational and experimental results were good, and demonstrated that the Euler code is capable of efficiently and accurately predicting the changes in the aerodynamic coefficients associated with inlet unstart and the effects of the nacelle and diverter components.

Description: This presentation will describe the organization and conduct of the workshops, list the topics discussed, and conclude with a more-detailed examination of a related set of issues dear to the presenters heart. Because the current HSCT configuration is expected to have (mostly) turbulent flow over the wings, and because current CFD predictions assume fully-turbulent flow, the wind tunnel testing to date has attempted to duplicate this condition at the lower Reynolds numbers attainable on the ground. This frequently requires some form of artificial boundary layer trip to induce transition near the wing's leading edge. But this innocent-sounding goal leads to a number of complications, and it is not clear that present-day testing technology is adequate to the task. An description of some of the difficulties, and work underway to address them, forms the "Results" section of this talk. Additional results of the testing workshop will be covered in presentations by other team members.

Description: This paper presents The Propulsion Airframe Integration Advisory report in viewgraph form. The approach of the advisory group is to identify and prioritize technology elements (1.0 Inlet Issues, 2.0 Nozzle Issues, 3.0 Nacelle Design, and 4.0 Airframe Integration).

Description: Preliminary human acceptability studies of sonic booms indicate that supersonic flight is unlikely to be acceptable even at noise levels significantly below 1994 low boom designs (reference 1, p. 288). Further, these low boom designs represent considerable changes to baseline configurations, and changes translate into additional effort and uncertain structural weight penalties that may provide no annoyance benefit, increasing the risk of including low boom technology. Since over land sonic boom designs were so risky (and yet the acceptability studies highlight how annoying sonic booms are), boom softening studies were undertaken to reduce the boom of baseline configurations using minor modifications that would not significantly change the designs. The goal of this work is to reduce boom levels over water. Even though Concorde over water boom has not been found to have any adverse environmental impact, boom levels for baseline HSCT designs are 50% higher in overpressure than the Concorde (due to a doubling in configuration weight with only a 50% increase in length),

Description: Officially, the Tu-144 was the first supersonic-cruise, passenger-carrying aircraft to enter commercial service. Design, construction, and testing were carried out by the Soviet Union, flight certification was by the Soviet Union, and the only regular passenger flights were scheduled and flown across the territory of the Soviet Union. Although it was not introduced to international passenger service, there were many significant engineering accomplishments achieved in the design, production, and flight of this aircraft. Development of the aircraft began with a prototype stage. Systematic testing and redesign led to a production aircraft in discrete stages that measurably improved the performance of the aircraft from the starting concept to final aircraft certification. It flew in competition with the English-French Concorde for a short time, but was withdrawn from national commercial service due to a lack of interest by airlines outside the Soviet Union. NASA became interested in the Tu- 144 aircraft when it was offered for use as a flying "testbed" in the study of operating characteristics of a supersonic-cruise commercial airplane. Since it had been in supersonic-cruise service, the Tu- 144 had operational characteris'tics similar to those anticipated in the conceptual aircraft designs being studied by the United States aircraft companies. In addition to the other operational tests being conducted on the Tu-144 aircraft, it was proposed that two sets of sonic-boom pressure signature measurements be made. The first set would be made on the ground, using techniques and devices similar to those in reference I and many other subsequent studies. A second set would be made in the air with an instrumented aircraft flying close under the Tu-144 in supersonic flight. Such in-flight measurements would require pressure gages that were capable of accurately recording the flow-field overpressures generated by the Tu- 144 at relatively close distances under the vehicle. Therefore, an analysis of the Tu-144 was made to obtain predictions of pressure signature shape and shock strengths at cruise conditions so that the range and characteristics of the required pressure gages could be determined well in advance of the tests. Cancellation of the sonic-boom signature measurement part of the tests removed the need for these pressure gages. Since CFD methods would be used to analyze the aerodynamic performance of the Tu-144 and make similar pressure signature predictions, the relatively quick and simple Whitham-theory pressure signature predictions presented in this paper could be used for comparisons. Pressure signature predictions of sonic-boom disturbances from the Tu- 144 aircraft were obtained from geometry derived from a three-view description of the production aircraft. The geometry was used to calculate aerodynamic performance characteristics at supersonic-cruise conditions. These characteristics and Whitham/Walkden sonic-boom theory were employed to obtain F-functions and flow-field pressure signature predictions at a Mach number of 2.2, at a cruise altitude of 61000 feet, and at a cruise weight of 350000 pounds.

Description: This document contains the details of the thermal analysis of the X-38 aft fin during re-entry. This analysis was performed in order to calculate temperature response of the aft fin components. This would be provided as input to a structural analysis and would also define the operating environment for the electromechanical actuator (EMA). The calculated structural temperature response would verify the performance of the thermal protection system (TPS). The geometric representation of the aft fin was derived from an I-DEAS finite element model that was used for structural analysis. The thermal mass network model was derived from the geometric representation.

Description: To summarize the significant highlights in this report: (1) Data quality, determined by multiple repeat runs performed on the TCA baseline configuration, and long-term repeatability, determined by comparing baseline Reference H data from this test to a previous test, have been shown to be good. (2) The longitudinal stability of the TCA is more non-linear than for the Reference H, and while it is similar at normal lift values, the TCA has considerably more pitch-up at higher lift. (3) Longitudinal control effectiveness of the TCA is similar to the Reference H and the ratio of elevator effectiveness to horizontal tail effectiveness is approximately 0.3. (4) The directional stability of the TCA is improved relative to Reference H at higher angles-of attack. The chine is effective for improving directional stability.

Description: The NASA-industry team has sponsored several studies in the last two years to address the installed nozzle boattail drag issues. Some early studies suggested that nozzle boattail drag could be as much as 25 to 40 percent of the subsonic cruise. As part of this study tests have been conducted at NASA-Langley to determine the uninstalled drag characteristics of a proposed nozzle. The overall objective was to determine the effects of nozzle external flap curvature and sidewall boattail variations. This test would also provide data for validating CFD predictions of nozzle boattail drag.

Description: AIRPLANE (Jameson/Baker) is a steady inviscid unstructured Euler flow solver. It has been validated on many HSR geometries. It is implemented as MESHPLANE, an unstructured mesh generator, and FLOPLANE, an iterative flow solver. The surface description from an Intergraph CAD system goes into MESHPLANE as collections of polygonal curves to generate the 3D mesh. The flow solver uses a multistage time stepping scheme with residual averaging to approach steady state, but R is not time accurate. The flow solver was ported from Cray to IBM SP2 by Wu-Sun Cheng (IBM); it could only be run on 4 CPUs at a time because of memory limitations. Meshes for the four cases had about 655,000 points in the flow field, about 3.9 million tetrahedra, about 77,500 points on the surface. The flow solver took about 23 wall seconds per iteration when using 4 CPUs. It took about eight and a half wall hours to run 1,300 iterations at a time (the queue limit is 10 hours). A revised version of FLOPLANE (Thomas) was used on up to 64 CPUs to finish up some calculations at the end. We had to turn on more communication when using more processors to eliminate noise that was contaminating the flow field; this added about 50% to the elapsed wall time per iteration when using 64 CPUs. This study involved computing lift and drag for a wing/body/nacelle configuration at Mach 0.9 and 4 degrees pitch. Four cases were considered, corresponding to four nacelle mass flow conditions.

Description: The primary objectives of this study were to expand the data base showing the effects of LE radius distribution and corresponding . sensitivity to Rn at subsonic and transonic conditions, and to assess the predictive capability of CFD for these effects. Several key elements led to the initiation of this project: 1) the necessity of meeting multipoint design requirements to enable a viable HSCT, 2) the demonstration that blunt supersonic leading-edges can be associated with performance gain at supersonic speeds , and 3) limited data. A test of a modified Reference H model with the TCA planform and 2 LE radius distributions was performed in the NTF, in addition to Navier-Stokes analysis for an additional 3 LE radius distributions. Results indicate that there is a tremendous potential to improve high-lift performance through the use of a blunt LE across the span given an integrated, fully optimized design, and that low Rn data alone is probably not sufficient to demonstrate the benefit.

Description: In cooperation with personnel from the Boeing ANP Laboratory and NASA Langley, a performance test was conducted using the Reference-H 1.675% model ("NASA Modular Model") without nacelles at the NASA Langley 16-Ft Transonic Tunnel. The main objective of the test was to determine the drag reduction achievable with leading-edge and trailing-edge flaps deflected along the outboard wing span at transonic Mach numbers (M = 0.9 to 1.2) for purpose of preliminary design and for comparison with computational predictions. The obtained drag data with flap deflections for Mach numbers of 1.07 to 1.20 are unique for the Reference H wing. Four leading-edge and two trailing-edge flap deflection angles were tested at a mean-wing chord-Reynolds number of about 5.7 million. An outboard-wing leading-edge flap deflection of 81 provides a 4.5 percent drag reduction at M = 1.2 A = 0.2), and much larger values at lower Mach numbers with larger flap deflections. The present results for the baseline (no flaps deflected) compare reasonably well with previous Boeing and NASA Ref-H tunnel tests, including high-Reynolds number NTF results. Viscous CFD simulations using the OVERFLOW thin-layer N.S. method properly predict the observed trend in drag reduction at M = 1.2 as function of leading-edge flap deflection. Modified linear theory properly predicts the flap effects on drag at subsonic conditions (Aero2S code), and properly predicts the absolute drag for the 40 and 80 leading-edge deflection at M = 1.2 (A389 code).

Description: Unstructured grid Euler computations, performed at supersonic cruise speed, are presented for a proposed high speed civil transport configuration, designated as the Technology Concept Airplane (TCA) within the High Speed Research (HSR) Program. The numerical results are obtained for the complete TCA cruise configuration which includes the wing, fuselage, empennage, diverters, and flow through nacelles at Mach 2.4 for a range of angles-of-attack and sideslip. The computed surface and off-surface flow characteristics are analyzed and the pressure coefficient contours on the wing lower surface are shown to correlate reasonably well with the available pressure sensitive paint results, particularly, for the complex shock wave structures around the nacelles. The predicted longitudinal and lateral/directional performance characteristics are shown to correlate very well with the measured data across the examined range of angles-of-attack and sideslip. The results from the present effort have been documented into a NASA Controlled-Distribution report which is being presently reviewed for publication.

Description: The objectives of the Cycle 2 Nonlinear Design Optimization Anlaytical Cross Checks are to: 1) Understand the variability in the predicted performance levels of the nonlinear designs arising from the use of different inviscid (full potential/Euler) and viscous (Navier-Stokes) analysis methods; and 2) Provide the information required to allow the performance levels of all three designs to be validated using the data from the NCV (nonlinear Cruise Validation) model test.

Description: During the last cycle of concept design and wind-tunnel testing, the goal of the low-boom- shaped HSCT concepts (the B-935, the LB-16, and the LB- 1 8) was to meet mission requirements and generate shaped, ground-level pressure signatures with nose shock strengths of 1.0 psf or less. The wind-tunnel tests of these concepts produced results that were partially successful and encouraging although not fully up to expectations. In spite of this, however, these conceptual designs were overly optimistic and not acceptable because: the wing planforms had excessive area; the wing structural aspect ratio was too high; one concept had aft-fuselage rather than under-the-wing engines; and the gross takeoff weights were unrealistically low because of engines that were early, high-tech versions of later, revised, more-realistic engines. The need for reducing the ground-level overpressure shock strengths still existed; a need to be met within more restrictive guidelines of mission performance and gross takeoff weight limitations. Therefore, it was decided that the next conceptual design cycle would focus on decreased nose shock strengths, "boom softening," in the signatures of the Boeing and the McDonnell Douglas baseline concepts rather than low-boom concepts with shaped-signature designs. Overly-optimistic results were not the only problem with these low-sonic-boom concepts. Papers given at the 1994 Sonic-Boom Workshop had demonstrated that the problem of successful nacelle integration on HSCT concepts had only been partially solved. Wind-tunnel pressure signature data, from the HSCT-11B (a.k.a. the LB-18) wind-tunnel model, showed that the Langley HSCT design and analysis method had been successful in reducing the nacelle-volume disturbances in the flow field. This was due.to the engine nacelles mounted behind the wing trailing-edge on the aft fuselage so that no nacelle-wing interference-lift flow-field disturbances were generated. While acceptable from a sonic-boom research point of view, this concept was unacceptable from several practical and structural considerations. Preliminary wind-tunnel pressure signature data from the LB-16 wind-tunnel model, which had the engine nacelles mounted under the wings (the usual location), indicated that the application of the Langley nacelle-integration method had been only partially successful in the reduction of the nacelle-volume with nacelle-wing interference-lift pressure disturbances. So, "boom softening" had to also address the task of successful integration of the engine nacelles, with the engines in the required under-the-wing location. Unless this problem was solved, low-sonic-boom and low-drag modifications to the wing planform, the airfoil shape, and the fuselage longitudinal area distribution could be nullified if the nacelle disturbances added increments to the nose-shock strengths that were removed through component tailoring. In this paper, an arrow-wing boom-softened HSC7 concept which incorporated modifications to a baseline McDonnell Douglas concept is discussed. The analysis of the concept's characteristics will include estimates of weight, center of gravity, takeoff field length, mission range, and predictions of its ground-level sonic-boom pressure signature. Additional modifications which enhanced the softened-boom performance of this concept are also described as well as estimates of the performance penalties induced by these modifications.

Description: A 1:300 scale wind-tunnel model of a conceptual High-Speed Civil Transport (HSCT) designed to generate a shaped, low-boom pressure signature on the ground was tested to obtain sonic-boom pressure signatures in the Langley Research Center Unitary Plan Wind Tunnel at a Mach number of 1.8 and a separation distance of about two body lengths or four wing-spans from the model. Two sets of engine nacelles representing two levels of engine technology were used on the model to determine the effects of increased nacelle volume. Pressure signatures were measured for (model lift)/(design lift) ratios of 0.5, 0.63, 0.75, and 1.0 so that the effect of lift on the pressure signature could be determined. The results of these tests were analyzed and used to discuss the agreement between experimental data and design expectations.

Description: The NASA High Speed Research (HSR) Program is intended to establish a technology base enabling industry development of an economically viable and environmentally acceptable second generation high speed civil transport (HSCT). The objective of the Configuration Aerodynamics task of the program is the development of aerodynamic drag reduction, stability and control, and propulsion airframe integration technologies required to support the HSCT development process. Aerodynamic design tools are being developed, evaluated, and validated through ground based experimental testing. In addition, methods for ground to flight scaling are being developed and refined.

Description: This paper reports on the model, test, and results from the Langley Supersonic Aftbody Closure wind tunnel test. This project is an experimental evaluation of the 1.5% Technology Concept Aircraft (TCA) aftbody closure model (Model 23) in the Langley Unitary Plan Wind Tunnel. The baseline TCA design is the result of a multidisciplinary, multipoint optimization process and was developed using linear design and analysis methods, supplemented with Euler and Navier-Stokes numerical methods. After a thorough design review, it was decided to use an upswept blade attached to the forebody as the mounting system. Structural concerns dictated that a wingtip support system would not be feasible. Only the aftbody part of the model is metric. The metric break was chosen to be at the fuselage station where prior aft-sting supported models had been truncated. Model 23 is thus a modified version of Model 20. The wing strongback, flap parts, and nacelles from Model 20 were used, whereas new aftbodies, a common forebody, and some new tails were fabricated. In summary, significant differences in longitudinal and direction stability and control characteristics between the ABF and ABB aftbody geometries were measured. Correcting the experimental data obtained for the TCA configuration with the flared aftbody to the representative of the baseline TCA closed aftbody will result in a significant reduction in longitudinal stability, a moderate reduction in stabilizer effectiveness and directional stability, and a moderate to significant reduction in rudder effectiveness. These reductions in the stability and control effectiveness levels of the baseline TCA closed aftbody are attributed to the reduction in carry-over area.

Description: There were two objectives for this test. First, was to assess the reasons why there is approximately 1.5 drag counts (cts) discrepancy between measured and computed drag improvement of the Non-linear Cruise Validation (NCV) over the Technology Concept Airplane (TCA) wing body (WB) configurations. The Navier-Stokes (N-S) pre-test predictions from Boeing Commercial Airplane Group (BCAG) show 4.5 drag cts of improvement for NCV over TCA at a lift coefficient (CL) of 0. I at Mach 2.4. The pre-test predictions from Boeing Phantom Works - Long Beach, BPW-LB, show 3.75 drag cts of improvement. BCAG used OVERFLOW and BPW-LB used CFL3D. The first test entry to validate the improvement was held at the NASA Langley Research Center (LARC) UPV;T, test number 1687. The experimental results showed that the drag improvement was only 2.6 cts, not accounting for laminar run and trip drag. This is approximately 1.5 cts less than predicted computationally. In addition to the low Reynolds Number (RN) test, there was a high RN test in the Boeing Supersonic Wind Tunnel (BSWT) of NCV and TCA. BSV@T test 647 showed that the drag improvement of NCV over TCA was also 2.6 cts, but this did account for laminar run and trip drag. Every effort needed to be done to assess if the improvement measured in LaRC UPWT and BSWT was correct. The second objective, once the first objective was met, was to assess the performance increment of NCV over TCA accounting for the associated laminar run and trip drag corrections in LaRC UPWT. We know that the configurations tested have laminar flow on portions of the wing and have trip drag due to the mechanisms used to force the flow to go from laminar to turbulent aft of the transition location.

Description: An improved laminar run and trip drag correction methodology for supersonic cruise performance testing was derived. This method required more careful analysis of the flow visualization images which revealed delayed transition particularly on the inboard upper surface, even for the largest trip disks. In addition, a new code was developed to estimate the laminar run correction. Once the data were corrected for laminar run, the correct approach to the analysis of the trip drag became evident. Although the data originally appeared confusing, the corrected data are consistent with previous results. Furthermore, the modified approach, which was described in this presentation, extends prior historical work by taking into account the delayed transition caused by the blunt leading edges.

Description: Gradient-based optimization requires accurate derivatives of the objective function and constraints. These gradients may have previously been obtained by manual differentiation of analysis codes, symbolic manipulators, finite-difference approximations, or existing automatic differentiation (AD) tools such as ADIFOR (Automatic Differentiation in FORTRAN). Each of these methods has certain deficiencies, particularly when applied to complex, coupled analyses with many design variables. Recently, a new AD tool called ADJIFOR (Automatic Adjoint Generation in FORTRAN), based upon ADIFOR, was developed and demonstrated. Whereas ADIFOR implements forward-mode (direct) differentiation throughout an analysis program to obtain exact derivatives via the chain rule of calculus, ADJIFOR implements the reverse-mode counterpart of the chain rule to obtain exact adjoint form derivatives from FORTRAN code. Automatically-generated adjoint versions of the widely-used CFL3D computational fluid dynamics (CFD) code and an algebraic wing grid generation code were obtained with just a few hours processing time using the ADJIFOR tool. The codes were verified for accuracy and were shown to compute the exact gradient of the wing lift-to-drag ratio, with respect to any number of shape parameters, in about the time required for 7 to 20 function evaluations. The codes have now been executed on various computers with typical memory and disk space for problems with up to 129 x 65 x 33 grid points, and for hundreds to thousands of independent variables. These adjoint codes are now used in a gradient-based aerodynamic shape optimization problem for a swept, tapered wing. For each design iteration, the optimization package constructs an approximate, linear optimization problem, based upon the current objective function, constraints, and gradient values. The optimizer subroutines are called within a design loop employing the approximate linear problem until an optimum shape is found, the design loop limit is reached, or no further design improvement is possible due to active design variable bounds and/or constraints. The resulting shape parameters are then used by the grid generation code to define a new wing surface and computational grid. The lift-to-drag ratio and its gradient are computed for the new design by the automatically-generated adjoint codes. Several optimization iterations may be required to find an optimum wing shape. Results from two sample cases will be discussed. The reader should note that this work primarily represents a demonstration of use of automatically- generated adjoint code within an aerodynamic shape optimization. As such, little significance is placed upon the actual optimization results, relative to the method for obtaining the results.

Description: Computations have been performed on the baseline Reference H wing/body configuration, as well as the Wing 704 configuration, an optimized wing and fuselage combination derived from Ref. H through automated optimization. The parabolized Navier-Stokes solver UPS was employed with viscous terms in two directions in an effort to understand the source and level of potential viscous/inviscid interactions. The paper briefly describes the UPS code and the grids used to obtain the solutions before the discussion of results. Results of these computations indicate that viscous/inviscid interaction can contribute increments to both the pressure- and friction-related drag. Computations were performed for wind tunnel conditions-1.675% scale models at a Reynolds number of 4 million per foot. Turbulent flow results were obtained using the Baldwin-Lomax algebraic turbulence model and were compared with laminar flow results. The laminar flow fields were used to obtain upper bounds on potential interaction effects.

Description: The NASA-industry team has sponsored several studies in the last two years to address the installed nozzle boattail drag issues. Some early studies suggested that nozzle boattail drag could be as much as 25 to 40 percent of the subsonic cruise. As part of this study tests have been conducted at NASA-Langley to determine the uninstalled drag characteristics of a proposed nozzle. The overall objective was to determine the effects of nozzle external flap curvature and sidewall boattail variations. This test would also provide data for validating CFD predictions of nozzle boattail drag.

Description: The objective of this milestone is to assess the propulsion/airframe integration characteristics of the Technology Concept Airplane and design variations through computational analysis and experimental subsonic through supersonic wind tunnel testing. The Milestone will generate a comprehensive CFD and wind tunnel data base of the baseline, and design variations. Emphasis will be placed on establishing the propulsion induced effects on the flight performance of the Technology Concept Airplane with all appropriate wind tunnel corrections.

Description: The computational fluid dynamics (CFD) comparisons being presented are compared to each other and to wind tunnel (WT) data on the baseline TCA. Some of the CFD computations were done prior to the tests and others later. Only force data (CL vs CD) from CFD will be presented as part of this report. The WT data presented comes from the testing of the baseline TCA in the Langley Unitary Plan Wind Tunnel (UPWT), Test Section #2. There are 2 sets of wind tunnel data being presented: one from test 1671 of model 2a (flapped wing) and the other from test 1679 of model 2b (solid wing). Most of the plots show only one run from each of the WT tests per configuration. But many repeat runs were taken during the tests. The WT repeat runs showed an uncertainty in the drag of +/- 0.5 count. There were times when the uncertainty in drag was better, +/- 0.25 count. Test 1671 data was of forces and pressures measured from model 2a. The wing had cutouts for installing various leading and trailing edge flaps at lower Mach numbers. The internal duct of the nacelles are not designed and fabricated as defined in the outer mold lines (OML) iges file. The internal duct was fabricated such that a linear transition occurs from the inlet to exhaust. Whereas, the iges definition has a constant area internal duct that quickly transitions from the inlet to exhaust cross sectional shape. The nacelle internal duct was fabricated, the way described, to save time and money. The variation in the cross sectional area is less than 1% from the iges definition. The nacelles were also installed with and without fairings. Fairings are defined as the build up of the nacelles on the upper wing surface so that the nacelles poke through the upper surface as defined in the OML iges file. Test 1679 data was of forces measured from model 2a and 2b. The wing for model 2b was a solid wing. The nacelles were built the same way as for model 2a, except for the nacelle base pressure installation. The nacelles were only tested with the fairings for model 2a and 2b during test 1679.

Description: Configuration design at Ames was carried out with the SYN87-SB (single block) Euler code using a 193 x 49 x 65 C-H grid. The Euler solver is coupled to the constrained (NPSOL) and the unconstrained (QNMDIF) optimization packages. Since the single block grid is able to model only wing-body configurations, the nacelle/diverter effects were included in the optimization process by SYN87's option to superimpose the nacelle/diverter interference pressures on the wing. These interference pressures were calculated using the AIRPLANE code. AIRPLANE is an Euler solver that uses a unstructured tetrahedral mesh and is capable of computations about arbitrary complete configurations. In addition, the buoyancy effects of the nacelle/diverters were also included in the design process by imposing the pressure field obtained during the design process onto the triangulated surfaces of the nacelle/diverter mesh generated by AIRPLANE. The interference pressures and nacelle buoyancy effects are added to the final forces after each flow field calculation. Full details of the (recently enhanced) ghost nacelle capability are given in a related talk. The pseudo nacelle corrections were greatly improved during this design cycle. During the Ref H and Cycle 1 design activities, the nacelles were only translated and pitched. In the cycle 2 design effort the nacelles can translate vertically, and pitch to accommodate the changes in the lower surface geometry. The diverter heights (between their leading and trailing edges) were modified during design as the shape of the lower wing changed, with the drag of the diverter changing accordingly. Both adjoint and finite difference gradients were used during optimization. The adjoint-based gradients were found to give good direction in the design space for configurations near the starting point, but as the design approached a minimum, the finite difference gradients were found to be more accurate. Use of finite difference gradients was limited by the CPU time limit available on the Cray machines. A typical optimization run using finite difference gradients can use only 30 to 40 design variables and one optimization iteration within the 8 hour queue limit for the chosen grid size and convergence level. The efficiency afforded by the adjoint method allowed for 50-120 design variables and 5-10 optimization iterations in the 8 hour queue. Geometric perturbations to the wing and fuselage were made using the Hicks/Henne (HH) shape functions. The HH functions were distributed uniformly along the chords of the wing defining sections and lofted linearly. During single-surface design, constraints on thickness and volume at selected wing stations were imposed. Both fuselage camber and cross-sectional area distributions were permitted to change during design. The major disadvantage to the use of these functions is the inherent surface waviness produced by repeated use of such functions. Many smoothing operations were required following optimization runs to produce a configuration with reasonable smoothness. Wagner functions were also used on the wing sections but were never used on the fuselage. The Wagner functions are a family of increasingly oscillatory functions that have also been used extensively in airfoil design. The leading and trailing edge regions of the wing were designed by use of polynomial and monomial functions respectively. Twist was attempted but was abandoned because of little performance improvement available from changing the baseline twist.

Description: The TetrUSS (Tetrahedral Unstructured Software System), developed at NASA LaRC, enables one to take a vehicle from its surface definition to its analyzed solution. The important parts are the shape definition, accomplished in GRIDTOOL; the initial front and volume grid generation in VGRID; the flow solver USM3D, and the various ways used to post-process the computational results.

Description: The objective of this study is to calibrate a Navier-Stokes code for the TCA (30/10) baseline configuration (partial span leading edge flaps were deflected at 30 degs. and all the trailing edge flaps were deflected at 10 degs). The computational results for several angles of attack are compared with experimental force, moments, and surface pressures. The code used in this study is CFL3D; mesh sequencing and multi-grid were used to full advantage to accelerate convergence. A multi-grid approach was used similar to that used for the Reference H configuration allowing point-to-point matching across all the trailingedge block interfaces. From past experiences with the Reference H (ie, good force, moment, and pressure comparisons were obtained), it was assumed that the mounting system would produce small effects; hence, it was not initially modeled. However, comparisons of lower surface pressures indicated the post mount significantly influenced the lower surface pressures, so the post geometry was inserted into the existing grid using Chimera (overset grids).

Description: The objective of the present study was to address the questions of: 1) how reliably or consistently the Navier-Stokes methods and processes used by the various organizations can predict integrated skin friction drag, and 2) how well the methods can predict trends within a family of optimized configurations. As a first step, all available skin friction drag predictions were accumulated to obtain a mean and standard deviation for the TCA (Technology Concept Airplane) baseline and each of the optimized configurations. It is observed that the optimization process has had little effect on the predicted skin friction drags. The variation in the mean that is observed is dwarfed by the standard deviations. In order to understand the reasons for the relatively large spreads in the computed results, a number of auxiliary computations have been performed using the UPS and OVERFLOW codes in an effort to identify and quantity potential sources of the variations.

Description: The paper presents the recent progress made towards developing an efficient and user-friendly parallel environment for routine analysis of large CFD problems. The coarse-grain parallel version of the CFL3D Euler/Navier-Stokes analysis code, CFL3Dhp, has been ported onto most available parallel platforms. The CFL3Dhp solution accuracy on these parallel platforms has been verified with the CFL3D sequential analyses. User-friendly pre- and post-processing tools that enable a seamless transfer from sequential to parallel processing have been written. Static load balancing tool for CFL3Dhp analysis has also been implemented for achieving good parallel efficiency. For large problems, load balancing efficiency as high as 95% can be achieved even when large number of processors are used. Linear scalability of the CFL3Dhp code with increasing number of processors has also been shown using a large installed transonic nozzle boattail analysis. To highlight the fast turn-around time of parallel processing, the TCA full configuration in sideslip Navier-Stokes drag polar at supersonic cruise has been obtained in a day. CFL3Dhp is currently being used as a production analysis tool.

Description: This paper presents an Unstructured Navier-Stokes Analysis of Full TCA (Technology Concept Airplane) Configuration. The topics include: 1) Motivation; 2) Milestone and approach; 3) Overview of the unstructured-grid system; 4) Results on full TCA W/B/N/D/E configuration; 5) Concluding remarks; and 6) Future directions.

Description: Automatic Grid Generation Wish List Geometry handling, including CAD clean up and mesh generation, remains a major bottleneck in the application of CFD methods. There is a pressing need for greater automation in several aspects of the geometry preparation in order to reduce set up time and eliminate user intervention as much as possible. Starting from the CAD representation of a configuration, there may be holes or overlapping surfaces which require an intensive effort to establish cleanly abutting surface patches, and collections of many patches may need to be combined for more efficient use of the geometrical representation. Obtaining an accurate and suitable body conforming grid with an adequate distribution of points throughout the flow-field, for the flow conditions of interest, is often the most time consuming task for complex CFD applications. There is a need for a clean unambiguous definition of the CAD geometry. Ideally this would be carried out automatically by smart CAD clean up software. One could also define a standard piece-wise smooth surface representation suitable for use by computational methods and then create software to translate between the various CAD descriptions and the standard representation. Surface meshing remains a time consuming, user intensive procedure. There is a need for automated surface meshing, requiring only minimal user intervention to define the overall density of mesh points. The surface mesher should produce well shaped elements (triangles or quadrilaterals) whose size is determined initially according to the surface curvature with a minimum size for flat pieces, and later refined by the user in other regions if necessary. Present techniques for volume meshing all require some degree of user intervention. There is a need for fully automated and reliable volume mesh generation. In addition, it should be possible to create both surface and volume meshes that meet guaranteed measures of mesh quality (e.g. minimum and maximum angle, stretching ratios, etc.).

Description: Conventional CFD methods and grids do not yield adequate resolution of the complex shock flow pattern generated by a real aircraft geometry. As a result, a unique grid topology and supersonic flow solver was developed at Northrop Grumman based on the characteristic behavior of supersonic wave patterns emanating from the aircraft. Using this approach, it was possible to compute flow fields with adequate resolution several body lengths below the aircraft. In this region, three-dimensional effects are diminished and conventional two-dimensional modified linear theory (MLT) can be applied to estimate ground pressure signatures or sonic booms. To accommodate real aircraft geometries and alleviate the burdensome grid generation task, an implicit marching multi-block, multi-grid finite-volume Euler code was developed as the basis for the sonic boom prediction methodology. The Thomas two-dimensional extrapolation method is built into the Euler code so that ground signatures can be obtained quickly and efficiently with minimum computational effort suitable to the aircraft design environment. The loudness levels of these signatures can then be determined using a NASA generated noise code. Since the Euler code is a three-dimensional flow field solver, the complete circumferential region below the aircraft is computed. The extrapolation of all this field data from a cylinder of constant radius leads to the definition of the entire boom corridor occurring directly below and off to the side of the aircraft's flight path yielding an estimate for the entire noise "annoyance" corridor in miles as well as its magnitude. An automated multidisciplinary sonic boom design optimization software system was developed during the latter part of HSR Phase 1. Using this system, it was found that sonic boom signatures could be reduced through optimization of a variety of geometric aircraft parameters. This system uses a gradient based nonlinear optimizer as the driver in conjunction with a computationally efficient Euler CFD solver (NIIM3DSB) for computing the three-dimensional near-field characteristics of the aircraft. The intent of the design system is to identify and optimize geometric design variables that have a beneficial impact on the ground sonic boom. The system uses a simple wave drag data format to specify the aircraft geometry. The geometry is internally enhanced and analytic methods are used to generate marching grids suitable for the multi-block Euler solver. The Thomas extrapolation method is integrated into this system, and hence, the aircraft's centerline ground sonic boom signature is also automatically computed for a specified cruise altitude and yields the parameters necessary to evaluate the design function. The entire design system has been automated since the gradient based optimization software requires many flow analyses in order to obtain the required sensitivity derivatives for each design variable in order to converge on an optimal solution. Hence, once the problem is defined which includes defining the objective function and geometric and aerodynamic constraints, the system will automatically regenerate the perturbed geometry, the necessary grids, the Euler solution, and finally the ground sonic boom signature at the request of the optimizer.

Description: A team was formed to tackle the sonic boom softening issues of the current Boeing HSCT design. The team consisted of personnel from NASA Ames, NASA Langley, and Boeing company. The work described in this paper was done when the first author was at NASA Ames Research Center. This paper presents the sonic boom softening work on two Boeing High Speed Civil Transport (HSCT) baseline configurations, Reference-H and Boeing-1122. This presentation can be divided into two parts: parametric studies and sonic boom minimization by CFD optimization routines.

Description: The objectives of this research are: 1) To determine the effect of geometric variations near the inboard leading-edge flap on high-lift and stability and control performance data; 2) To determine Re effects on TCA (Technology Concept Aircraft) high-lift configuration for optimum high-lift and stability and control performance at takeoff, climbout, approach and landing conditions; and 3) To obtain flow-visualization data on upper surface of wing for CFD validations. This paper is presented in viewgraph form.

Description: This presentation describes the advances being made with the Aerodynamic Shape Optimization (ASO) and high-fidelity Multidisciplinary Optimization (MDO) software used in the High Speed Research Program at NASA Ames Research Center. The description starts with the motivation for continued ASO/MDO development. Objectives of the current work are then presented. A list of ingredients deemed necessary for a flexible design environment is discussed, and the HSR requirement for different geometries at different design points is explained. Multiple design disciplines within a high-fidelity design environment are demonstrated. Finally, progress so far is summarized and planned future work is outlined.

Description: This report considers the effect of canard and horizontal tail vertical position on the aerodynamic characteristics of the PTC configuration without nacelles and diverters. This analysis is followed by three optimization studies using canard and tail incidence as design variables in the first problem followed by an optimization run with canard and tail incidence and wing camber design variables and finally an optimization run with canard incidence and wing camber. The first problem was run at fixed lift while the other two problems were run at fixed angle of attack. The final investigation reported here will show data from a component buildup study using the PTC configuration. This final study will show the aerodynamic interference between the canard, wing and horizontal tail.

Description: This paper presents results of three minor studies into the behavior of the OVERFLOW with respect to the prediction of skin friction drag on wing bodies at cruise Mach number and wind tunnel Reynolds number. The studies include a preliminary assessment of the behavior of the two new 2-equation turbulence models introduced with the latest version of OVERFLOW (v. 1.8f), an investigation into potential improvements in the matrix dissipation scheme currently implemented in OVERFLOW, and an analysis of the observed sensitivity of the code's skin friction predictions to grid stretching at solid surface boundaries.

Description: The computational results of the optimized complete configurations, including nacelles and diverters, are presented in terms of drag count improvement compared with the TCA baseline configuration at Mach 2.4, C(sub L)=0.1. The three candidate designs are designated by the organization from which they were derived. ARC represents the Ames Research Center 1-03 design, BCAG represents the Boeing Commercial Aircraft Group's design from Seattle, and BLB represents the design from Boeing Long Beach. All CFD methods are in unanimous agreement that the Ames 1-03 configuration has the largest performance improvement, followed closely by the BCAG configuration, with a much smaller improvement attained by Boeing Long Beach. The Ames design was obtained using the single-block wing/body code SYN87-SB with its "pseudo" nacelle option-an elaborate technique for incorporating nacelle/diverter effects into the design optimization process. This technique uses AIRPLANE surface pressure coefficient data with and without the nacelles/diverters. Further details of this method are described. It is reasonable to expect that further improvements could be achieved by including the "real" nacelles directly into the optimization process by use of the newly-developed multiblock optimization code, SYN107-MB, which can handle full configurations.

Description: The aim of this work is to demonstrate a simple technique which accounts for aeroelastic deformations experienced by HSR wind-tunnel models within CFD computations. With improved correlations, CFD can become a more effective tool for augmenting the post-test understanding of experimental data. The present technique involves the loose coupling of a low-level structural representation within the ELAPS code, to an unstructured Navier-Stokes flow solver, USM3Dns. The ELAPS model is initially calibrated against bending characteristics of the wind-tunnel model. The strength of this method is that, with a single point calibration of a simple structural representation, the static aeroelastic effects can be accounted for in CFD calculations across a range of test conditions. No prior knowledge of the model deformation during the wind-on test is required. This approach has been successfully applied to the high aspect-ratio planforms of subsonic transports. The current challenge is to adapt the procedure to low aspect-ratio planforms typical of HSR configurations.

Description: This presentation includes three topics: (1) Analysis of isolated boattail drag; (2) Computation of Technology Concept Airplane (TCA)-installed nacelle effects on aerodynamic performance; and (3) Assessment of TCA inlet flow quality.

Description: LaRC conducted a code validation study for the OVERFLOW code to ascertain its accuracy for boattail drag prediction. The OVERFLOW results compared favorably with the LaRC 16-ft. Transonic Wind Tunnel (TWT) data, and prior CFD solutions from PAB3D and CFL3D. The ultimate goal is to investigate the installation drag of the nacelle boattails with powered nozzles at transonic mach numbers. The OVERFLOW solver was chosen because of its ability to accept volume overlapping structured grid for very complex airframe configurations. Structured grid components for representing the transonic nozzle boattail can be added to the BCAG grid for a TCA airframe with 2D bifurcated inlet and flow through nacelle without alteration. The focus of this research was to determine the suitability of the OVERFLOW solver for accomplishing this ultimate goal. This presentation will first introduce the transonic nozzle boattail wind-tunnel model geometry, followed by an examination of aerodynamic features based on the current OVERFLOW solutions and the solutions obtained previously using PAB3D, comparisons of Cp on the flap surface between the OVERFLOW solutions, wind tunnel data, and solutions from other CFD codes, an assessment of boattail drag count prediction, and a work plan for FY99.

Description: A concept of system identification applied to high performance aircraft is introduced followed by a discussion on the identification methodology. Special emphasis is given to model postulation using time invariant and time dependent aerodynamic parameters, model structure determination and parameter estimation using ordinary least squares and mixed estimation methods. At the same time problems of data collinearity detection and its assessment are discussed. These parts of methodology are demonstrated in examples using flight data of the X-29A and X-31A aircraft. In the third example wind tunnel oscillatory data of the F-16XL model are used. A strong dependence of these data on frequency led to the development of models with unsteady aerodynamic terms in the form of indicial functions. The paper is completed by concluding remarks.

Description: This paper reports the predicted M = 2.4 strut-interference effects on a closed aftbody with empennage for the TCA baseline model. The strut mounting technique was needed in order to assess the impact of aft-end shaping, i.e. open for a sting or closed to better represent a flight vehicle. However,this technique can potentially lead to unanticipated effects that are measured on the aft body. Therefore, a set of computations were performed in order to examine the closed aft body with and without strut present, at both zero and non-zero angles of sideslip (AOS). The work was divided into a computational task performed by Javier A. Garriz, using an inviscid (Euler) solver, and a monitoring/reporting task done by John E. Lamar. All this work was performed during FY98 at the NASA Langley Research Center.

Description: This CFD experiment concludes that the potential difference between the flow between a flight Reynolds number test and a sub-scale wind tunnel test are substantial for this particular nozzle boattail geometry. The early study was performed using a linear k-epsilon turbulence model. The present study was performed using the Girimaji formulation of a algebraic Reynolds stress turbulent simulation.

Description: The Lockheed Martin spillage study was a substantial effort and is worthy of a separate paper. However, since a paper was not submitted a few of the most pertinent results have been pulled out and included in this paper. The reader is urged to obtain a copy of the complete Boeing Configuration Aerodynamics final 1995 contract report for the complete Lockheed documentation of the spillage work. The supersonic cruise studies presented here focus on the bifurcated - axisymmetric inlet drag delta. In the process of analyzing this delta several test/CFD data correlation problems arose that lead to a correction of the measured drag delta from 4.6 counts to 3.1 counts. This study also lead to much better understanding of the OVERFLOW gridding and solution process, and to increased accuracy of the force and moment data. Detailed observations of the CFD results lead to the conclusion that the 3.1 count difference between the two inlet types could be reduced to approximately 2 counts, with an absolute lower bound of 1.2 counts due to friction drag and the bifurcated lip bevel.

Description: The Configuration Aerodynamics (CA) element of the High Speed Research (HSR) program is managed by a joint NASA and Industry team, referred to as the Technology Integration Development (ITD) team. This team is responsible for the development of a broad range of technologies for improved aerodynamic performance and stability and control characteristics at subsonic to supersonic flight conditions. These objectives are pursued through the aggressive use of advanced experimental test techniques and state of the art computational methods. As the HSR program matures and transitions into the next phase the objectives of the Configuration Aerodynamics ITD are being refined to address the drag reduction needs and stability and control requirements of High Speed Civil Transport (HSCT) aircraft. In addition, the experimental and computational tools are being refined and improved to meet these challenges. The presentation will review the work performed within the Configuration Aerodynamics element in 1994 and 1995 and then discuss the plans for the 1996-1998 time period. The final portion of the presentation will review several observations of the HSR program and the design activity within Configuration Aerodynamics.

Description: The NASA High-Speed Research program developed the High-Lift Engine Aeroacoustics Technology (HEAT) program to demonstrate satisfactory interaction between the jet noise suppressor and high-lift system of a High-Speed Civil Transport (HSCT) configuration at takeoff, climb, approach and landing conditions. One scheme for reducing jet exhaust noise generated by an HSCT is the use of a mixer-ejector system which would entrain large quantities of ambient air into the nozzle exhaust flow through secondary inlets in order to cool and slow the jet exhaust before it exits the nozzle. The effectiveness of such a noise suppression device must be evaluated in the presence of an HSCT wing high-lift system before definitive assessments can be made concerning its acoustic performance. In addition, these noise suppressors must provide the required acoustic attenuation while not degrading the thrust efficiency of the propulsion system or the aerodynamic performance of the high-lift devices on the wing. Therefore, the main objective of the HEAT program is to demonstrate these technologies and understand their interactions on a large-scale HSCT model. The HEAT program is a collaborative effort between NASA-Ames, Boeing Commercial Airplane Group, Douglas Aircraft Corp., Lockheed-Georgia, General Electric and NASA - Lewis. The suppressor nozzles used in the tests were Generation 1 2-D mixer-ejector nozzles made by General Electric. The model used was a 13.5%-scale semi-span model of a Boeing Reference H configuration.

Description: The mission of High-Lift Technology is to develop technology allowing the design of practical high lift concepts for the High-Speed Civil Transport (HSCT) in order to: 1) operate safely and efficiently; and 2) reduce terminal control area and community noise. In fulfilling this mission, close and continuous coordination will be maintained with other High-Speed Research (HSR) technology elements in order to support optimization of the overall airplane (rather than just the high lift system).

Description: Nozzle boattail drag is significant for the High Speed Civil Transport (HSCT) and can be as high as 25% of the overall propulsion system thrust at transonic conditions. Thus, nozzle boattail drag has the potential to create a thrust-drag pinch and can reduce HSCT aircraft aerodynamic efficiencies at transonic operating conditions. In order to accurately predict HSCT performance, it is imperative that nozzle boattail drag be accurately predicted. Previous methods to predict HSCT nozzle boattail drag were suspect in the transonic regime. In addition, previous prediction methods were unable to account for complex nozzle geometry and were not flexible enough for engine cycle trade studies. A computational fluid dynamics (CFD) effort was conducted by NASA and McDonnell Douglas to evaluate the magnitude and characteristics of HSCT nozzle boattail drag at transonic conditions. A team of engineers used various CFD codes and provided consistent, accurate boattail drag coefficient predictions for a family of HSCT nozzle configurations. The CFD results were incorporated into a nozzle drag database that encompassed the entire HSCT flight regime and provided the basis for an accurate and flexible prediction methodology.

Description: High-lift system performance will have a large impact on airplane noise and weight. Successful completion of PCD1 activities provided greater understanding of aerodynamic characteristics and configuration features important to high-lift system performance including: 1) Reynolds number effects (Ref. H); 2) Propulsion/airframe integration effects; and 3) Planform effects, canard/3-surface, alternate high-lift concepts, etc. PCD2 plans are aimed at achieving technology development performance goals and increasing technology readiness level for Technology Concept.

Description: Experiments were conducted in the NASA Ames 9-Ft by 7-Ft Supersonic and 11-Ft by 11-Ft Transonic Wind Tunnels of a 2.7% Reference H (Ref. H) Nacelle Airframe Interference (NAI) High Speed Civil Transport (HSCT) model. NASA Ames did the experiment with the cooperation and assistance of Boeing and McDonnell Douglas. The Ref. H geometry was designed by Boeing. The model was built and tested by NASA under a license agreement with Boeing. Detailed forces and pressures of individual components of the configuration were obtained to assess nacelle airframe interference through the transonic and supersonic flight regime. The test apparatus was capable of measuring forces and pressures of the Wing body (WB) and nacelles. Axisymmetric and 2-D inlet nacelles were tested with the WB in both the in-proximity and captive mode. The in-proximity nacelles were mounted to a nacelle support system apparatus and were individually positioned. The right hand nacelles were force instrumented with flow through strain-gauged balances and the left hand nacelles were pressure instrumented. Mass flow ratio was varied to get steady state inlet unstart data. In addition, supersonic spillage data was taken by testing the 2-D inlet nacelles with ramps and the axisymmetric inlet nacelles with an inlet centerbody for the Mach condition of interest. The captive nacelles, both axisymmetric and 2-D, were attached to the WB via diverters. The captive 2-D inlet nacelle was also tested with ramps to get supersonic spillage data. Boeing analyzed the data and showed a drag penalty of four drag counts for the 2-D compared with the axisymmetric inlet nacelle. Two of the four counts were attributable to the external bevel designed into the 2-D inlet contour. Boeing and McDonnell Douglas used these data for evaluating Computational Fluid Dynamic (CFD) codes and for evaluation of nacelle airframe integration problems and solutions.

Description: This paper discusses a method for the identification and application of reduced-order models based on linear and nonlinear aerodynamic impulse responses. The Volterra theory of nonlinear systems and an appropriate kernel identification technique are described. Insight into the nature of kernels is provided by applying the method to the nonlinear Riccati equation in a non-aerodynamic application. The method is then applied to a nonlinear aerodynamic model of an RAE 2822 supercritical airfoil undergoing plunge motions using the CFL3D Navier-Stokes flow solver with the Spalart-Allmaras turbulence model. Results demonstrate the computational efficiency of the technique.

Description: The objective of this study is to calibrate a Navier-Stokes code for a high-lift Reference H configuration using structured grids. The outline of this presentation will first include a brief description of the grids used and the flow solver. Next the results will be presented in terms of convergence and resources used on the C-90. Predicted force and moment and surface pressure results are compared to experiment and off- and on-surface flow viz. is discussed.

Description: With the continuous progress in hardware and numerical schemes, Computational Unsteady Aerodynamics (CUA), that is, the application of Computational Fluid Dynamics (CFD) to unsteady flowfields, is slowly finding its way as a useful and reliable tool (turbulence and transition modeling permitting) in the aircraft, helicopter, engine and missile design and development process. Before a specific code may be used with confidence it is essential to validate its capability to describe the physics of the flow correctly, or at least to the level of approximation required, for which purpose a comparison with accurate experimental data is needed. Unsteady wind tunnel testing is difficult and expensive; two factors which dramatically limit the number of organizations with the capability and/or resources to perform it. Thus, unsteady experimental data is scarce, often classified and scattered in diverse documents. Additionally, access to the reports does not necessarily assure access to the data itself. The collaborative effort described in this paper was conceived with the aim of collecting into a single easily accessible document as much quality data as possible. The idea is not new. In the early 80's NATO's AGARD (Advisory Group for Aerospace Research & Development) Structures and Material Panel (SMP) produced AGARD Report No. 702 "Compendium of Unsteady Aerodynamic Measurements", which has found and continues to find extensive use within the CUA Community. In 1995 AGARD's Fluid Dynamics Panel (FDP) decided to update and expand the former database with new geometries and physical phenomena, and launched Working Group WG-22 on "Validation Data for Computational Unsteady Aerodynamic Codes". Shortly afterwards AGARD was reorganized as the RTO (Research and Technology Organization) and the WG was renamed as AVT (Applied Vehicle Technolology) WG-003. Contributions were received from AEDC, BAe, DLR, DERA, Glasgow University, IAR, NAL, NASA, NLR, and ONERA. The final publication with the results of the exercise is expected in the second part of 1999. The aim of the present paper is to announce and present the new database to the Aeroelasticity community. It is also intended to identify, together with one of the groups of end users it targets, deficiencies in the compendium that should be addressed by means of new wind tunnel tests or by obtaining access to additionally existing data.

Description: The HSR program moved into phase two with the selection of a new airplane configuration, the Technology Concept Airplane (TCA). The TCA was designed based on the experiences gained while investigating both the Reference H and the Arrow Wing configurations in different wind tunnels and CFD studies. Part of that investigation included performing extensive high Reynolds number testing on the Reference H configuration in the NTF to provide data for predicting full-scale flight performance, as well as developing techniques for testing these types of configurations in the NTF. With the selection of the TCA configuration, a smaller investigation was designed to examine whether or not the scaling characteristics of the TCA configuration are similar to those observed for the Reference H configuration. This presentation will include a description of the 2.2% Modified Reference H model used in this investigation (highlighting the similarities and the differences when compared to the TCA configuration), the testing objectives, and some preliminary findings that are relevant to the current high-lift system.

Description: The codes WINGDES2 and AERO2S were easy to obtain, and technical help was readily available. The codes have a long, well-documented history of successful optimizations of various aircraft configurations. The codes were easy to use, although specification of input data was time-consuming. The Run times were short, allowing the many runs necessary for the Suction Parameter matrix to be accomplished within a day or two. The results of the optimization appear to be reasonable.

Description: Full-scale flight performance predictions can be made using CFD or a combination of CFD and analytical skin-friction predictions. However, no matter what method is used to obtain full-scale flight performance predictions knowledge of the boundary layer state is critical. The implementation of CFD codes solving the Navier-Stokes equations to obtain these predictions is still a time consuming, expensive process. In addition, to ultimately obtain accurate performance predictions the transition location must be fixed in the CFD model. An example, using the M2.4-7A geometry, of the change in Navier-Stokes solution with changes in transition and in turbulence model will be shown. Oil flow visualization using the M2.4-7A 4.0% scale model in the 14'x22' wind tunnel shows that fixing transition at 10% x/c in the CFD model best captures the flow physics of the wing flow field. A less costly method of obtaining full-scale performance predictions is the use of non-linear Euler codes or linear CFD codes, such as panel methods, combined with analytical skin-friction predictions. Again, knowledge of the boundary layer state is critical to the accurate determination of full-scale flight performance. Boundary layer transition detection has been performed at 0.3 and 0.9 Mach numbers over an extensive Reynolds number range using the 2.2% scale Reference H model in the NTF. A temperature sensitive paint system was used to determine the boundary layer state for these conditions. Data was obtained for three configurations: the baseline, undeflected flaps configuration; the transonic cruise configuration; and, the high-lift configuration. It was determined that at low Reynolds number conditions, in the 8 to 10 million Reynolds number range, the baseline configuration has extensive regions of laminar flow, in fact significantly more than analytical skin-friction methods predict. This configuration is fully turbulent at about 30 million Reynolds number for both 0.3 and 0.9, Mach numbers. Both the transonic cruise and the high-lift configurations were fully turbulent aft of the leading-edge flap hingeline at all Reynolds numbers.

Description: The Transonic Nozzle Boattail Drag Study was initiated in 1995 to develop an understanding of how external nozzle transonic aerodynamics effect airplane performance and how strongly those effects are dependent on nozzle configuration (2D vs. axisymmetric). MDC analyzed the axisymmetric nozzle. Boeing subcontracted Northrop-Grumman to analyze the 2D nozzle. AU participants analyzed the AGARD nozzle as a check-out and validation case. Once the codes were checked out and the gridding resolution necessary for modeling the separated flow in this region determined, the analysis moved to the installed wing/body/nacelle/diverter cases. The boat tail drag validation case was the AGARD B.4 rectangular nozzle. This test case offered both test data and previous CFD analyses for comparison. Results were obtained for test cases B.4.1 (M=0.6) and B.4.2 (M=0.938) and compared very well with the experimental data. Once the validation was complete a CFD grid was constructed for the full Ref. H configuration (wing/body/nacelle/diverter) using a combination of patched and overlapped (Chimera) grids. This was done to ensure that the grid topologies and density would be adequate for the full model. The use of overlapped grids allowed the same grids from the full configuration model to be used for the wing/body alone cases, thus eliminating the risk of grid differences affecting the determination of the installation effects. Once the full configuration model was run and deemed to be suitable the nacelle/diverter grids were removed and the wing/body analysis performed. Reference H wing/body results were completed for M=0.9 (a=0.0, 2.0, 4.0, 6.0 and 8.0), M=1.1 (a=4.0 and 6.0) and M=2.4 (a=0.0, 2.0, 4.4, 6.0 and 8.0). Comparisons of the M=0.9 and M=2.4 cases were made with available wind tunnel data and overall comparisons were good. The axi-inlet/2D nozzle nacelle was analyzed isolated. The isolated nacelle data coupled with the wing/body result enabled the interference effects of the installed nacelles to be determined. Isolated nacelle mm were made at M=0.9 and M=1.1 for both the supersonic and transonic nozzle settings. AU of the isolated nacelle cases were run at alpha=0. Full configuration runs were to be made at Mach numbers of 0.9, 1.1, and 2.4 (the same as the wing/body and isolated nacelles). Both the isolated nacelles and installed nacelles were run with inlet conditions designed to give zero spillage. This was to be done in order to isolate the boattail effects as much as possible. Full configuration runs with the supersonic nozzles were completed for M=0.9 and 1.1 at a=4.0 and 6.0 (4 runs total) and with the transonic nozzles at M=0.9 and 1.1 at a=2.0, 4.0 and 6.0 (6 runs total). Drag breakdowns were completed for the M=0.9 and M= 1.1 showing favorable interference drag for both cases.

Description: This paper describes an Overview of Technology Integration. The topics include: 1) Technology Concept Airplane Description; 2) LCAP Overview; and 3) ACE (Aeroelastic Concept Engineering) Overview.

Description: The SLFC (Supersonic Laminar Flow Control) mission is to develop and validate technologies and perform the SLFC aerodynamic design for the HSCT with an assessment of the net benefits and risks.

Description: This paper describes HSR (High Speed Research) Aerodynamic Performance Status and Challenges. The topics include: 1) Aero impact on HSR; 2) Goals and Targets; 3) Progress and Status; and 4) Remaining Challenges. This paper is presented in viewgraph form.

Description: The Propulsion Induced Effects (PIE) test program is being lead by NASA Ames for Configuration Aerodynamics (CA). Representatives from CA, Technology Integration (TI), Inlet, and the Nozzle ITD's are working with Ames in defining and executing this test program. The objective of the CA 4-14 milestone is to assess the propulsion/airframe integration characteristics of the Technology Concept Airplane (TCA) and design variations using computational and experimental methods. The experimental aspect includes static calibrations, transonic and supersonic wind tunnel testing. The test program will generate a comprehensive database that will include all appropriate wind tunnel corrections, with emphasis placed on establishing the propulsion induced effects on the flight performance of the TCA.

Description: A synoptic of programmatic and technical lessons learned in the development of aerothermodynamic data bases for the X-33 and X-34 programs is presented in general terms and from the perspective of the NASA Langley Research Center Aerothermodynamics Branch. The format used is that of the aerothermodynamic chain, the links of which are personnel, facilities, models/test articles, instrumentation, test techniques, and computational fluid dynamics (CFD). Because the aerodynamic data bases upon which the X-33 and X-34 vehicles will fly are almost exclusively from wind tunnel testing, as opposed to CFD, the primary focus of the lessons learned is on ground-based testing.

Description: Two CD-ROMs contain experimental ice shapes and code prediction used for validation of LEWICE 2.0 (see NASA/CR-1999-208690, CASI ID 19990021235). The data include ice shapes for both experiment and for LEWICE, all of the input and output files for the LEWICE cases, JPG files of all plots generated, an electronic copy of the text of the validation report, and a Microsoft Excel(R) spreadsheet containing all of the quantitative measurements taken. The LEWICE source code and executable are not contained on the discs.

Description: The objective of this study was to investigate compressibility effects on a high-lift flowfield by simulating the flow about a three-dimensional multi-element wing. The computations were performed by solving both the incompressible and compressible Navier-Stokes equations (using the INS3D and OVERFLOW codes) on structured, overset grids. Turbulence was modeled via the one-equation, fully turbulent Spalart-Allmaras model. The computational results were validated with surface pressure measurements acquired at the NASA Ames 7- by 10-Foot Wind Tunnel. The geometry used for all computations consisted of an unswept wing in a landing configuration with a half-span flap and a three-quarter-span slat mounted inside a rectangular duct approximating the wind tunnel walls. The solutions were carefully examined to account for effects due to differences in algorithms. Compressibility effects were demonstrated by comparing surface particle traces, sectional pressure coefficient and boundary layer profile plots. It was found that small regions of compressibility near the slat and main-element leading edge can largely impact the flow. Even small compressibility regions can have significant global effects on the circulation and separation of each of the high-lift elements.

Description: A historical overview will be presented of the research conducted on the structure and modification of the vortices generated by the lifting surfaces of subsonic transport aircraft. The seminar will describe the three areas of vortex research; namely, the magnitude of the hazard posed, efforts to reduce the hazard to an acceptable level, and efforts to develop a systematic means for avoiding vortex wakes. It is first pointed out that the characteristics of lift-generated vortices are related to the aerodynamic shapes that produce them and that various arrangements of surfaces can be used to produce different vortex structures. The largest portion of the research conducted to date has been directed at finding ways to reduce the hazard potential of lift-generated vortices shed by subsonic transport aircraft in the vicinity of airports during landing and takeoff operations. It is stressed that lift-generated vortex wakes are so complex that progress towards a solution requires application of a combined theoretical and experimental research program because either alone often leads to incorrect conclusions. It is concluded that a satisfactory aerodynamic solution to the wake-vortex problem at airports has not yet been found but a reduction in the impact of the wake-vortex hazard on airport capacity may become available in the foreseeable future through wake-vortex avoidance concepts currently under study. The material to be presented in this overview is drawn from articles published in aerospace journals that are available publicly.

Description: The effects of periodic excitation on the evolution of a turbulent jet were studied experimentally. A short, wide-angle diffuser was attached to the jet exit and excitation was introduced at the junction between the jet exit and the diffuser inlet. The introduction of high amplitude periodic excitation at the jet exit enhances the mixing and promotes attachment of the jet shear-layer to the diffuser wall. Vectoring is achieved by applying the excitation over a fraction of the circumference of the circular jet, enhancing its spreading rate on the excited side and its tendency to reattach to that side. Static deflection studies demonstrate that the presence of the wide-angle diffuser increases the effectiveness of the added periodic momentum due to a favorable interaction between the excitation, the jet shear-layer and the diffuser wall. This point was further demonstrated by the evolution of a wave packet that was excited in the jet shear-layer. Strong amplification of the wave packet was measured with a diffuser attached to the jet exit. The turbulent jet responds quickly (10-20 msec) to step changes in the level of the excitation input. The response scales with the jet exit velocity and is independent of the Reynolds number. Jet deflection angles were found to be highly sensitive to the relative direction between the excitation and the jet flow and less sensitive to the excitation frequency. The higher jet deflection angles were obtained for a diffuser length of about two diameters and for diffusers with half-angles greater than 15 degrees.

Description: The effects of ambient turbulence on decay and descent of aircraft wake vortices are studied using a validated, three-dimensional: large-eddy simulation model. Numerical simulations are performed in order to isolate the effect of ambient turbulence on the wake vortex decay rate within a neutrally-stratified atmosphere. Simulations are conducted for a range of turbulence intensities, by injecting wake vortex pairs into an approximately homogeneous and isotropic turbulence field. The decay rate of the vortex circulation increases clearly with increasing ambient turbulence level, which is consistent with field observations. Based on the results from the numerical simulations, simple decay models are proposed as functions of dimensionless ambient turbulence intensity (eta) and dimensionless time (T) for the circulation averaged over a range of radial distances. With good agreement with the numerical results, a Gaussian type of vortex decay model is proposed for weak turbulence: while an exponential type of Tortex decay model can be applied for strong turbulence. A relationship for the vortex descent based on above vortex decay model is also proposed. Although the proposed models are based on simulations assuming neutral stratification, the model predictions are compared to Lidar vortex measurements observed during stable, neutral, and unstable atmospheric conditions. In the neutral and unstable atmosphere, the model predictions appear to be in reasonable agreement with the observational data, while in the stably-stratified atmosphere, they largely underestimate the observed circulation decay with consistent overestimation of the observed vortex descent. The underestimation of vortex decay during stably-stratified conditions suggests that stratification has an important influence on vortex decay when ambient levels of turbulence are weak.

Description: A sensitivity study for the in-ground effect on aircraft wake vortices has been conducted using a validated large eddy simulation model. The numerical results are compared with observed data and show good agreement for vortex decay and lateral vortex transport. The vortex decay rate is strongly influenced by the ground, but appears somewhat insensitive to ambient turbulence. In addition, the results show that the ground can affect the trajectory and descent-rate of a wake vortex pair at elevations up to about 3 b(sub o) (where b(sub o) is the initial vortex separation). However, the ground does not influence the average circulation of the vortices until the cores descend to within about 0.6 b(sub o), after which time the ground greatly enhances their rate of demise. Vortex rebound occurs in the simulations, but is more subtle than shown in previous numerical studies.

Description: The benchmark active controls technology and wind tunnel test program at NASA Langley Research Center was started with the objective to investigate the nonlinear, unsteady aerodynamics and active flutter suppression of wings in transonic flow. The paper will present the flutter suppression control law design process, numerical nonlinear simulation and wind tunnel test results for the NACA 0012 benchmark active control wing model. The flutter suppression control law design processes using classical, and minimax techniques are described. A unified general formulation and solution for the minimax approach, based on the steady state differential game theory is presented. Design considerations for improving the control law robustness and digital implementation are outlined. It was shown that simple control laws when properly designed based on physical principles, can suppress flutter with limited control power even in the presence of transonic shocks and flow separation. In wind tunnel tests in air and heavy gas medium, the closed-loop flutter dynamic pressure was increased to the tunnel upper limit of 200 psf. The control law robustness and performance predictions were verified in highly nonlinear flow conditions, gain and phase perturbations, and spoiler deployment. A non-design plunge instability condition was also successfully suppressed.

Description: A review of compressibility effects on dynamic stall of pitching airfoils and unsteady separation control by manipulation of unsteady vorticity using a deformable leading edge airfoil design is presented.

Description: This program has involved, first of all, a critical state-of-the-art assessment of vortex-body interactions. Then, efforts were focused on experimental investigation on coupled-wake instabilities and turbulence occurring in a two-cylinder system. An extensive review was undertaken on the effect of incident vortices on various types of bodies. These incident vortices have a length scale of the same order of magnitude as the scale of the body. The body can take on various forms, including, for example, a circular cylinder, a blade or a wing. The classes of vortex-body interaction that were critically assessed include: (1) Periodic distortion of the incident (primary) vortex and shedding of secondary vorticity from the surface of the body. (2) Modulated vortex distortion and shedding at a leading-edge or surface due to incidence of a complex system of vortices. (3) Vortex distortion and shedding in presence of body oscillation. (4) Three-dimensional vortex interaction and shedding. For all of these classes of vortex-body interaction, quantitative topologies of the vorticity distributions and streamline patterns were found to be central to a unified description of mechanisms of vortex distortion and shedding. In most cases, it was possible to define relationships between vortex interactions and unsteady loading at the body surface. This phase of the program was an experimental investigation of a two-cylinder system, which simulated a central aspect of a four-wheel bogie on a large-scale commercial aircraft. The overall aim of this experimental research program was to determine the crucial elements of the unsteadiness in the gap and near-wake regions as a function of time using cinema-based techniques. During the research program, various image evaluation techniques were employed. They involved assessment of instantaneous velocity fields, streamline topology and patterns of vorticity. Experiments were performed in a large-scale water channel using a high-resolution version of digital particle image velocimetry. The program has focused on acquisition of images of velocity and vorticity for varying gap widths between the two-cylinder system. As a result of analysis of a relatively large number of images, it is demonstrated that low frequency instabilities can occur in the gap region between the cylinder. These low frequency instabilities are hypothesized to influence the near-wake structure of the entire two-cylinder system. The nature of the unstable shear layers in the gap region involves generation of small-scale Kelvin-Helmholtz instabilities. These unsteady shear layers then impinge upon the upper and lower surfaces of the cylinders, thereby influencing both the unsteady structure and the time-averaged patterns of the near-wake. Initial efforts have focused on characterization of the patterns of instantaneous and averaged streamlines using topological concepts. The end result of this investigation is a series of documented instantaneous images. They will serve as a basis for various types of post-processing, which will lead to a fuller understanding of the instantaneous and time-averaged unstable-turbulent fields in the gap region and downstream of the two-cylinder system. This further assessment is the focus of a subsequent program.

Description: DARWIN is a web-based system for presenting the results of wind-tunnel testing and computational model analyses to aerospace designers. DARWIN captures the data, maintains the information, and manages derived knowledge (e.g. visualizations, etc.) of large quantities of aerospace data. In addition, it provides tools and an environment for distributed collaborative engineering. We are currently constructing the third version of the DARWIN software system. DARWN's development history has, in some sense, tracked the development of web applications. The 1995 DARWIN reflected the latest web technologies--CGI scripts, Java applets and a three-layer architecture--available at that time. The 1997 version of DARWIN expanded on this base, making extensive use of a plethora of web technologies, including Java/JavaScript and Dynamic HTML. While more powerful, this multiplicity has proven to be a maintenance and development headache. The year 2000 version of DARWIN will provide a more stable and uniform foundation environment, composed primarily of Java mechanisms. In this paper, we discuss this evolution, comparing the strengths and weaknesses of the various architectural approaches and describing the lessons learned about building complex web applications.

Description: A design optimization procedure for improved sonic boom and aerodynamic performance of high speed aircraft is presented. The multiobjective optimization procedure simultaneously minimizes the sonic boom at a given distance from the aircraft and the drag-to-lift ratio (C(sub D)/C(sub L)) Of the aircraft. Upper and lower bounds are also imposed on the lift coefficient. The Kreisselmeier - Steinhauser function is used for the multiobjective optimization formulation. A discrete semi-analytical aerodynamic sensitivity analysis procedure coupled with an analytical grid sensitivity analysis technique is used for evaluating design sensitivities. The use of the semi-analytical sensitivity analysis techniques results in significant computational savings. The flow equations are solved using a three-dimensional parabolized Navier-Stokes solver. Sonic boom analysis is performed using an extrapolation procedure. A nonlinear programming technique and an approximate analysis procedure are used for the optimization. The optimization procedure developed is applied to the design of two high speed configurations, namely, a doubly swept wing-body configuration and a delta wing-body configuration. For the two sweep case only, minimization of the first peak in the pressure signature is performed first by optimizing only the nose radius and length of the aircraft. Minimization of the second peak in the pressure signature is performed next by optimizing only the wing geometric parameters. Significant improvements are obtained in the sonic boom characteristics and the aerodynamic performance of the wing-body configurations.

Description: A study of the Trailblazer vehicle inlet was conducted using the Global Air Sampling Program (GASP) code for flight Mach numbers ranging from 4-12. Both perfect gas and finite rate chemical analysis were performed with the intention of making detailed comparisons between the two results. Inlet performance was assessed using total pressure recovery and kinetic energy efficiency. These assessments were based upon a one-dimensional stream-thrust-average of the axisymmetric flowfield. Flow visualization utilized to examine the detailed shock structures internal to this mixed-compression inlet. Kinetic energy efficiency appeared to be the least sensitive to differences between the perfect gas and finite rate chemistry results. Total pressure recovery appeared to be the most sensitive discriminator between the perfect gas and finite rate chemistry results for flight Mach numbers above Mach 6. Adiabatic wall temperature was consistently overpredicted by the perfect gas model for flight Mach numbers above Mach 4. The predicted shock structures were noticeably different for Mach numbers from 6-12. At Mach 4, the perfect gas and finite rate chemistry models collapse to the same result.

Description: The primary objective of this paper is to demonstrate that the field of aeroelasticity continues to play a critical role in the design of modern aerospace vehicles, and several important problems are still far from being well understood. Furthermore, the emergence of new technologies, such as the use of adaptive materials (sometimes denoted as smart structures technology), providing new actuator and sensor capabilities, has invigorated aeroelasticity, and generated a host of new and challenging research topics that can have a major impact on the design of a new generation of aerospace vehicles.

Description: Compressible and incompressible versions of a three-dimensional unstructured mesh Reynolds-averaged Navier-Stokes flow solver have been differentiated and resulting derivatives have been verified by comparisons with finite differences and a complex-variable approach. In this implementation, the turbulence model is fully coupled with the flow equations in order to achieve this consistency. The accuracy demonstrated in the current work represents the first time that such an approach has been successfully implemented. The accuracy of a number of simplifying approximations to the linearizations of the residual have been examined. A first-order approximation to the dependent variables in both the adjoint and design equations has been investigated. The effects of a "frozen" eddy viscosity and the ramifications of neglecting some mesh sensitivity terms were also examined. It has been found that none of the approximations yielded derivatives of acceptable accuracy and were often of incorrect sign. However, numerical experiments indicate that an incomplete convergence of the adjoint system often yield sufficiently accurate derivatives, thereby significantly lowering the time required for computing sensitivity information. The convergence rate of the adjoint solver relative to the flow solver has been examined. Inviscid adjoint solutions typically require one to four times the cost of a flow solution, while for turbulent adjoint computations, this ratio can reach as high as eight to ten. Numerical experiments have shown that the adjoint solver can stall before converging the solution to machine accuracy, particularly for viscous cases. A possible remedy for this phenomenon would be to include the complete higher-order linearization in the preconditioning step, or to employ a simple form of mesh sequencing to obtain better approximations to the solution through the use of coarser meshes. An efficient surface parameterization based on a free-form deformation technique has been utilized and the resulting codes have been integrated with an optimization package. Lastly, sample optimizations have been shown for inviscid and turbulent flow over an ONERA M6 wing. Drag reductions have been demonstrated by reducing shock strengths across the span of the wing. In order for large scale optimization to become routine, the benefits of parallel architectures should be exploited. Although the flow solver has been parallelized using compiler directives. The parallel efficiency is under 50 percent. Clearly, parallel versions of the codes will have an immediate impact on the ability to design realistic configurations on fine meshes, and this effort is currently underway.

Description: Transition to turbulence in swept-wing flows has resisted correlation with linear theory because of its sensitivity to freestream conditions and 3-D roughness and because one of the principal instability modes quickly 'becomes nonlinear. In the face of such a formidable problem, two rather long-term fundamental efforts have been underway at DLR Gottinberg and Arizona State University that address swept-wing transition. These efforts have been recently reviewed by Bippes (1997) and Reibert and Saric (1997). Thus, the present work is a continuation of a series of studies on swept-wing boundary layers which have led to a better understanding of the transition process. In particular, we have taken advantage of the sensitivity to 3-D roughness and the modal nature of the instability in order to propose a particular control strategy. Complementing the two aforementioned reviews, general reviews of the swept-wing transition problem are found in Arnal (1997) and Kachanov (1996). Other recent reviews include Reshotko (t997), Crouch (1997), and Herbert (1997a,b). The failure of linear theory is discussed in Reed et al. (1996). The historical work is found in Reed and Sar-ic (1989). The basic idea is that the combination of sweep and chordwise pressure gradient within the boundary layer creates a velocity component perpendicular to the inviscid streamline. This crossflow profile is inflectional and exhibits both traveling and stationary unstable waves called crossflow vortices that are (approximately) aligned along the inviscid streamlines. Under conditions of low freestream turbulence levels, the dominant crossflow wave is stationary (Reibert and Saric t997) while moderate to high turbulence levels initiate dominant traveling waves (Dehle and Bippes 1996; Bippes 1997). 'Me mechanism is relatively insensitive to sound and 2-D surface roughness (Radeztsky et al. 1993) but very sensitive to 3-D roughness near the attachment line. We concentrate our work on low-turbulence freestream flows and stationary crossflow waves. Although the v' and w' components of the disturbances are very small, by convecting streamwise momentum in the wall-normal direction, they produce 0(l) changes inu'.Thus the mean flow is highly distorted with localized inflection points. Transition is then triggered by a high-frequency secondary instability of the distorted mean profile.

Description: The concept of "numerical speed of sound" is proposed in the construction of numerical flux. It is shown that this variable is responsible for the accurate resolution of' discontinuities, such as contacts and shocks. Moreover, this concept can he readily extended to deal with low speed and multiphase flows. As a results, the numerical dissipation for low speed flows is scaled with the local fluid speed, rather than the sound speed. Hence, the accuracy is enhanced the correct solution recovered, and the convergence rate improved. We also emphasize the role of mass flux and analyze the behavior of this flux. Study of mass flux is important because the numerical diffusivity introduced in it can be identified. In addition, it is the term common to all conservation equations. We show calculated results for a wide variety of flows to validate the effectiveness of using the numerical speed of sound concept in constructing the numerical flux. We especially aim at achieving these two goals: (1) improving accuracy and (2) gaining convergence rates for all speed ranges. We find that while the performance at high speed range is maintained, the flux now has the capability of performing well even with the low: speed flows. Thanks to the new numerical speed of sound, the convergence is even enhanced for the flows outside of the low speed range. To realize the usefulness of the proposed method in engineering problems, we have also performed calculations for complex 3D turbulent flows and the results are in excellent agreement with data.

Description: This paper examines flight-measured subsonic lift and drag characteristics of seven lifting-body and wing-body reentry vehicle configurations with truncated bases. The seven vehicles are the full-scale M2-F1, M2-F2, HL-10, X-24A, X-24B, and X-15 vehicles and the Space Shuttle prototype. Lift and drag data of the various vehicles are assembled under aerodynamic performance parameters and presented in several analytical and graphical formats. These formats unify the data and allow a greater understanding than studying the vehicles individually allows. Lift-curve slope data are studied with respect to aspect ratio and related to generic wind-tunnel model data and to theory for low-aspect-ratio planforms. The proper definition of reference area was critical for understanding and comparing the lift data. The drag components studied include minimum drag coefficient, lift-related drag, maximum lift-to-drag ratio, and, where available, base pressure coefficients. The effects of fineness ratio on forebody drag were also considered. The influence of forebody drag on afterbody (base) drag at low lift is shown to be related to Hoerner's compilation for body, airfoil, nacelle, and canopy drag. These analyses are intended to provide a useful analytical framework with which to compare and evaluate new vehicle configurations of the same generic family.

Description: A sailplane is currently being developed at NASA's Dryden Flight Research Center to support a high altitude flight experiment. The purpose of the experiment is to measure the performance characteristics of an airfoil at altitudes between 100,000 and 70,000 feet at Mach numbers between 0.65 and 0.5. The airfoil lift and drag are measured from pilot and static pressures. The location of the separation bubble and vortex shedding are measured from a hot film strip. The details of the flight experiment are presented. A comparison of several estimates of the airfoil performance is also presented. The airfoil, APEX-16, was designed by Drela (MIT) with his MSES code. A two dimensional Navier-Stokes analysis has been performed by Tatineni and Zhong (UCLA) and another at the Dryden Flight Research Center. The role these analysis served to define the experiment is discussed.

Description: Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

Description: A recently developed neural net-based aerodynamic design procedure is used in the redesign of a transonic turbine stage to improve its unsteady aerodynamic performance. The redesign procedure used incorporates the advantages of both traditional response surface methodology and neural networks by employing a strategy called parameter-based partitioning of the design space. Starting from the reference design, a sequence of response surfaces based on both neural networks and polynomial fits are constructed to traverse the design space in search of an optimal solution that exhibits improved unsteady performance. The procedure combines the power of neural networks and the economy of low-order polynomials (in terms of number of simulations required and network training requirements). A time-accurate, two-dimensional, Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the optimization procedure. The procedure yielded a modified design that improves the aerodynamic performance through small changes to the reference design geometry. These results demonstrate the capabilities of the neural net-based design procedure, and also show the advantages of including high-fidelity unsteady simulations that capture the relevant flow physics in the design optimization process.

Description: An investigation was performed to evaluate leading-and trailing-edge flap deflections for optimal aerodynamic performance of a High-Speed Civil Transport concept during takeoff and approach-to-landing conditions. The configuration used for this study was designed by the Douglas Aircraft Company during the 1970's. A 0.1-scale model of this configuration was tested in the Langley 30- by 60-Foot Tunnel with both the original leading-edge flap system and a new leading-edge flap system, which was designed with modem computational flow analysis and optimization tools. Leading-and trailing-edge flap deflections were generated for the original and modified leading-edge flap systems with the computational flow analysis and optimization tools. Although wind tunnel data indicated improvements in aerodynamic performance for the analytically derived flap deflections for both leading-edge flap systems, perturbations of the analytically derived leading-edge flap deflections yielded significant additional improvements in aerodynamic performance. In addition to the aerodynamic performance optimization testing, stability and control data were also obtained. An evaluation of the crosswind landing capability of the aircraft configuration revealed that insufficient lateral control existed as a result of high levels of lateral stability. Deflection of the leading-and trailing-edge flaps improved the crosswind landing capability of the vehicle considerably; however, additional improvements are required.

Description: The third High-Speed Research Sonic Boom Workshop was held at NASA Langley Research Center on June 1-3, 1994. The purpose of this workshop was to provide a forum for Government, industry, and university participants to present and discuss progress in their research. The workshop was organized into sessions dealing with atmospheric propagation; acceptability studies; and configuration design, and testing. Attendance at the workshop was by invitation only. The workshop proceedings include papers on design, analysis, and testing of low-boom high-speed civil transport configurations and experimental techniques for measuring sonic booms. Significant progress is noted in these areas in the time since the previous workshop a year earlier. The papers include preliminary results of sonic boom wind tunnel tests conducted during 1993 and 1994 on several low-boom designs. Results of a mission performance analysis of all low-boom designs are also included. Two experimental methods for measuring near-field signatures of airplanes in flight are reported.

Description: The High-Speed Research Program sponsored the NASA High-Speed Research Program Aerodynamic Performance Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of: Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization) and High-Lift. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. The HSR AP Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas within the airframe element of the HSR Program. This Volume 2/Part 1 publication presents the High-Lift Configuration Development session.