Integrated Computational/Experimental Approach to ...

4th Symposium on Integrating CFD and Experiments in Aerodynamics 14-16 September 2009 von Karman Institute, Rhode-Saint-Genèse, Belgium

Integrated Computational/Experimental Approach to UCAV and Delta-Canard Configurations Regarding Stability & Control Andreas Schütte1 Russell M. Cummings2 Thomas Loeser3 Dan D. Vicroy4 1

German Aerospace Center (DLR), Institute of Aerodynamics and Flow Technology Lilienthalplatz 7, 38108 Braunschweig, Germany e-mail: [email protected] 2 United States Air Force Academy (USAFA), Department of Aeronautics USAF Academy, CO 80840, USA e-mail: [email protected] 3 German-Dutch Wind Tunnels (DNW), Low-Speed-Wind-Tunnel Braunschweig (NWB) Lilienthalplatz 7, 38108 Braunschweig, Germany e-mail: [email protected] 4 NASA Langley Research Center (LaRC) Hampton, VA 23681-2199, USA e-mail: [email protected]

Abstract In the past computational fluid dynamics (CFD) was often relegated to the position of following experimental investigations of aerodynamic vehicles. As CFD has matured and developed, the place of CFD in the vehicle development process has improved substantially, however a true integration of CFD with experiments has been slow to develop. This paper outlines the integrated approach to simulating static and dynamic stability characteristics for a generic UCAV and the X-31 configuration, as being performed by NATO RTO Task Group AVT-161. The UCAV, named SACCON (Stability And Control CONfiguration), and the X-31 are the subject of an intensive computational and experimental study. The stability characteristics of the vehicles are being evaluated via a highly integrated approach, where CFD and experimental results are being used in a parallel and collaborative fashion. The purpose of the paper is to show how these investigations have been performed in a truly integrated fashion, which has noticeably improved the knowledge gained from the program. Key words: CFD, experiment, aerodynamics, Stabilty&Control, UCAV, X-31, RTO. α β η Θ Ψ V f t k

= = = = = = = = =

s y/s y x/cr x z cref cr cp

AoA, Angle of attack [°] AoS, Angle of side slip [°] Flap deflection angle [°] Pitch angle [°] Yaw angle [°] On flow velocity [m/s] Frequency [Hz] Time [s] Reduced frequency 2π·f·cref /V

1

= = = = = = = = =

Half span [m] Non-dimensional span wise coordinate Span wise coordinate [m] Non-dimensional chord wise Chord wise coordinate [m] Vertical coordinate [m] Reference length [m] Length of root chord [m] Pressure coefficient


DoF CFRP PIV LE SLE RLE RLE-FT MRP VKK AVT

= = = = = = = = = =

Degree of Freedom Carbon Fiber Reinforced Plastic Particle Image Velocimetry Leading Edge Sharp Leading Edge Round Leading Edge Round Leading Edge – Fixed Transition Moment Reference Point Leading Edge Flap Applied Vehicle Technology

cL cm cl cn cS cp cz cY cmx cmy cmz

= = = = = = = = = = =

Lift coefficient [-] Pitch moment coefficient (AE) [-] Roll moment coefficient (AE) [-] Yaw moment coefficient (AE) [-] Side force coefficient (AE) [-] Pressure coefficient [-] Normal force coefficient (MF) [-] Side force coefficient (MF) [-] Roll moment coefficient (MF) [-] Pitch moment coefficient (MF) [-] Yaw moment coefficient (MF) [-]

Introduction Traditionally, aerodynamics research has been conducted with a “Serial Approach”, with either initial theoretical or empirical predictions being followed by experimental testing, or in some cases the experiments were run alone. During the early development of CFD, comparisons with experiment were necessary to develop confidence in algorithms and methods, so CFD researchers were often left to sort through experimental data, and often wonder about the details of how the tests were conducted since the original researchers may not have been available to answer questions (see Figure 1). Comparisons were done well after the experiments, and often suffered from a lack of detailed information about the experiment (such as boundary conditions, detailed free stream conditions, precise geometry information, etc.). As time progressed, and as improvements were made to the capabilities of CFD, CFD researchers asked for experiments to be Experiment CFD performed specifically for Long time gap (years or decades) validation and verification of their predictions. These experiments were often completed prior to any CFD Figure 1: Early “Serial Approach”; 1960s through 1980s. analysis being conducted, and the CFD researchers were then allowed to use the data. While the time gap shown in Fig. 1 had been eliminated, there was still no (or little) interaction between experimental and CFD researchers. Eventually, as CFD became more accepted during the late 1980s and 1990s, CFD was included in the research process, as shown in CFD Prediction Experiment CFD Verification Figure 2. This “Alternating Serial interaction, but limited (or no) integration Serial Approach” had the advantage that the experimental and computational researchers were somewhat Figure 2: “Alternating Serial Approach”; late 1980s through present. interacting and integrating their work, which meant that the results of the research were hopefully improved. Unfortunately, the work could still be done with a mentality that did not fully integrate the approach, since the CFD researchers could “finish” their initial estimates and then hand off their work to the experimentalists, and the experimentalists could do the same thing once they were done with the tests. If researchers truly wanted to integrate their work, it would only happen if they chose to become involved in the work being done by the experimental research group. As the benefits of working in an integrated fashion became more and more apparent over the past decade or so, some researchers were able to work in a more integrated “parallel” fashion (see Figure 3). This meant that CFD predictions, which began prior to the experimentation in order to inform the researchers how best to use their

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limited experimental resources, could now continue in parallel with the experiments. This was done in order to help answer questions that arose during Experiment the experiments and also to provide the experimentalists Parallel interaction with limited integration with detailed flow field results while the experiFigure 3: “Early Parallel Approach”; late 1990s through present. ments were going on. The CFD researchers would then use the experimental results as they became available in order to verify and validate their earlier predictions, as well as to answer any lingering questions from the experimental campaign. While this approach was much superior to the earlier approaches, the interaction between the computational and experimental researchers was still somewhat limited, and did not take full advantage of the expertise of each group. The purpose of this paper is to outline how a NATO Research and Technology Organization (RTO) Task Group (AVT-161) has attempted to learn from the experiences of the past and to improve the integration of CFD and experiments within the realms of a state-of-the-art research program. Specifically, we will outline the overall goals of the Task Group and discuss how we have approached our research in order to take advantage of the strengths of both computations and experiments. Details of our experiences will be discussed and recommendations for future improvements will be made. CFD Prediction

CFD Verification

AVT-161 Task Group The RTO AVT-161 Task Group was established to improve the ability to accurately predict both static and dynamic stability of air and sea vehicles using computational fluid dynamics (CFD) methods. Whereas this paper will concentrate on the air vehicle application within the task group, the idea is to identify major synergy in terms of physical modeling, fluid structures or transition effects. The task group identified three major topics of interest: 1) the experimental group to provide highly accurate static and dynamic validation data, 2) the CFD community group who Contribution was trying to predict the steady state and dynamic Configuration SACCON Organisation Exp. CFD Engineering S&C configuration behavior of the target Methods X-31 SACCON +model configurations, and 3) the NASA x x x x x Stability and Control DLR x x x x x x (S&C) group which is USAFA x x x x analyzing the experiDNW x x x mental and numerical EADS-MAS x x x data. The objective of the BAE Systems x x x x task group is to provide Dstl x x best practice procedures DSTO x x x x to predict the static and Boeing x x dynamic behavior esFOI x x pecially for configuraKTH x x x tions with vortex Metacomp Techn. x x x dominated flow fields NLR x x x where non-linear effects Nangia ARA x x x have a significant impact. NEAR x x x These nonlinear regimes ONERA x x x x are the areas where usual QinetiQ x linear S&C methods fail Univ. of Liverpool x x x or wind tunnel data are Univ. of Quebec x x x only available for non TU Braunschweig Vortical flow consultant flight flow regimes. NRC Vortical flow consultant Currently these deficienTable 1: AVT-161 Task Group participants and contributions. cies can only be

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addressed through costly flight testing. Because of this the main focus is the prediction with CFD methods than enhancing existing S&C system identification methods. The AVT-161 partners and their contributions are listed in Table 1. Background Both AVT-161 target configurations possess a delta wing plan form with medium-swept leading edges between 45° and 57°. The leading edge nose radii vary from sharp to medium and large roundness. That implies that all flow features of common UCAV and fighter aircraft configurations and aerodynamic challenging problems have to be solved or captured by the use of current computational methods and models. The X-31 configuration was used as a target configuration within the DLR project Sikma – “Simulation of complex maneuvers” [1]-[3]. Within the DLR project several experimental and numerical investigations were established. Based on these investigations, additional wind tunnel tests were done by DLR to complete the data set and knowledge of the wind tunnel model characteristics. These results were then available for use by the AVT-161 Task Group. Another set of important research on delta wing configurations with rounded leading edges had been developed within a previous Task Group, AVT-113 [4][5]. In AVT-113 the focus was on experimental and numerical investigations on delta wing configurations with various leading edges from sharp to various round radii. The research started with fundamental wind tunnel data obtained by NASA [6], and was followed by several pre-test CFD results which supported the follow-on wind tunnel investigations, which use various advanced experimental methods. All of these investigations resulted in a highly improved understanding of the flow physics, as well as meeting the objective to achieve best practice methods for computational simulation. One further result was to identify the weaknesses of computational simulation and areas of interest related to the flow physics of delta wings with round leading edges. The results of the AVT-113 task group have been presented at the AIAA Aerospace Sciences Meeting in Reno during 2008, with special sessions devoted to both experimental [7]-[11] and numerical [12]-[19] work; many of these papers have and will appear in the Journal of Aircraft. The second target configuration of the AVT-113 task group was the F-16XL CAWAPI configuration. The results of this part of the task group which is more related to the X-31 configuration was published in the special section of the Journal of Aircraft by Lamar and Obara [20][21], Boelens et al. [22][23], Görtz et al. [24] and Fritz et al. [25]. A summery of lessons learned is also given by Rizzi et al. [26]. These previous investigations directly led to the creation of AVT-161, and the lessons learned and best practices of AVT-113 served as the foundation to the formation of AVT-161. In addition to the work of the NATO RTO Task Groups, there have been several other investigations focused on the prediction of delta wing configurations in maneuvering flight. Recent experimental and CFD investigations of the dynamic behavior of a pitching UCAV configuration was presented by Cummings et al. [27], as well as investigations on time-dependant flow predictions [28]. The dynamic aerodynamic effects on the vortical flow of pitching and rolling delta wings are discussed in several works, including Ericsson [29] and Arthur et al.[30], whereas experimental investigations are given by Chaderjian and Schiff [31][32], Hanff et al. [33], Hummel and Löser [34]. Within a DARPA project several experimental investigations on aerodynamic structural mechanics coupling effects have been analyzed by wind tunnel tests using a UCAV model configuration, as reported by Kudva and Sanders et al. [35][36]. Another important issue which is relevant to delta wing configurations in maneuvering flight is the extent of the laminar flow regions. A recent contribution to this topic is given by Arthur et al. [37]. The CFD prediction of the laminar/turbulent transition of a UCAV configuration is discussed in this paper. Finally, experimental and numerical aerodynamic behavior for a UCAV has been investigated for the 1303 UCAV configuration. The 1303 configuration was established by US AFRL, who offered the configuration for study to other partners, as well as for the UK within “The Technical Cooperation Program” (TTCP). The 1303 configuration is a tail-less “flying-wing” with a leading-edge sweep angle of 47°. Other related configurations have been investigated, and their results are reports in publications by Zhang et al. [38], Petterson [39][40], Wong [41], Arthur [42] and Sherer et al. [43]. CFD/Experiment Integration Since many of the participants in AVT-161 had been associated with AVT-113, a great deal of experience had been obtained prior to starting detailed work on the research for AVT-161. Specifically, we had learned that conducting any of the serial approaches to “integrated” research (as shown in Fig. 1 and 2), did not best utilize

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the capabilities of the researchers on the Task Group, and probably did not lead to the best possible use of time (especially for a program with a limited life, which most programs have in today’s world). Because of these experienCFD Verification ces, we undertook to insure CFD Prediction CFD Analysis and Design that an integrated approach to research took place, and that as much interaction between computational and Experiment 2 Experiment 1 experimental researchers took place throughout the Parallel interaction with higher‐order integration program. We wanted to insure that our research was Figure 4: “Alternating Parallel Approach”; improved approach. done more like the “Alternating Parallel Approach” shown in Figure 4, where CFD was used for early predictions of the configuration (here I am primarily discussing the SACCON configuration, since the X-31 had been designed and tested prior to the start of AVT-161). In addition, the computational researchers provided the experimentalists with a detailed list of the type of data they would like to see collected in order to maximize the benefit of the CFD analysis: 1.

2.

3. 4.

Have at least two to three tunnel entries (possibly in multiple tunnels) so that the following can be accomplished: a. measure static aerodynamics first (at various angles of attack and sideslip, with and without control surface deflections) so that the CFD community can begin to understand the details of the flow and have a chance to validate our predictions and aid in defining the parameters for the follow-on tunnel entries b. measure dynamic derivatives and/or perform dynamic maneuvers—there was a large variation of ideas here, but we can certainly discuss this in Athens; the model scale needs to be appropriate for realistic reduced frequencies and for multiple wind tunnels It is essential to accurately know and verify the tunnel conditions, including: a. freestream conditions, including flow angularity and variability b. tunnel wall effects c. model deformation d. model position in the tunnel It is also essential to know the experimental uncertainty for all measured quantities for both static and dynamic testing. Important results should be repeated to aid in this knowledge. Besides forces/moments and surface pressures, the following flow visualization suggestions would help the CFD community a great deal: a. knowledge of flow near the leading edge, trailing edge, and wing tips b. unsteady surface measurements, including surface pressures and flow patterns (accomplished with pressure transducers, surface oil flows, and/or PSP) c. knowledge about transition locations d. unsteady flow field measurements, including knowledge about vortex development and interactions via off-surface flow visualization (accomplished with PIV and/or LDA)

The over-arching theme of the requests from the prediction community can best be summarized as: “understand the developing flow structures.” In other words, the prediction community not only needs to know the gross aerodynamics of the vehicle, but also the causes of any interesting or unusual flow phenomenon. This request was nearly unanimous and strongly stated, even though the researchers realized the request would be difficult for most wind tunnels to fully obtain. The early CFD predictions were used to design the wind tunnel model and were quickly followed by the first phase of experimental testing. Once the initial results were available, then a phase of verification and further design could take place using CFD, which would quickly be followed by a second experimental campaign. Finally, CFD could be used to perform final analysis and explanation for the experimental results, and increased knowledge of the ability to predict the S&C of aircraft would result.

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Specifically, the SACCON configuration research proRedesign ject has taken place along the timeline shown in “Wish List” provided Initial CFD & lowerFigure 5. Initial planning for to experimentalists order simulation Predictions the Task Group was taking (CFD and Low-Order) Static CFD of redesign Static & dynamic CFD place during 2007, resulting Lower-order modeling in a proposal of the SACCON configuration in Data analysis Fall of 2007 by DLR and Test planning S&C analysis Experiments EADS-MAS. That initial design was examined with DNW-NWB Entry 1 both lower-order modeling DNW-NWB Entry 2 and CFD during the last half of 2007, during which time NASA Langley 14X22 it was determined that the Timeline 2007 2008 2009 2010 design could be improved with some relatively minor Figure 5: Our Approach: Integrated Design/Compute/Test Timeline. re-design changes, which took place during the first two months of 2008 with inputs from EADS-MAS, Nangia Consulting, BAE Systems, and DLR. Once the redesigned configuration was finalized, detailed test planning began by DLR and German-Dutch Wind Tunnels (DNW) for two entries in the DNW-NWB wind tunnel. This was accompanied in parallel by numerous static CFD simulations being conducted by DLR, EADS-MAS, NASA Langley, and the US Air Force Academy. Other CFD static simulations were completed prior to the end of wind tunnel testing by the University of Liverpool, FOI, DSTL, DSTO, NLR, Metacomp, and ONERA. Lower-order model estimates of the static aerodynamics were also conducted by Nangia Consulting and NEAR. The ability of these various approaches to predict the static and dynamic behavior of the SACCON configuration forms the foundation of the purpose of the Task Group. The first campaign in the DNW-NWB tunnel included force and moment testing of both static and dynamic cases. Once this data was available a number of organizations began performing dynamic simulations using CFD including NLR, NASA Langley, and the US Air Force Academy. The simulations were being conducted at the same time the second campaign took place in the DNW-NWB tunnel, which was primarily to perform flow visualization studies. PIV measurements were made by both DLR and ONERA, and this off-surface flow information will form a crucial part of understanding the flow field and resulting forces and moments of SACCON. In addition, initial analysis of the stability and control characteristics of SACCON has already started and will continue in parallel with the CFD simulations and the additional wind tunnel test being planned for the end of 2009 at NASA Langley in the 14′×22′ wind tunnel. The timeline in Fig. 5 shows how an integrated approach to a fairly complex design/compute/test program can be conducted in a relatively compressed time schedule. The progress made to date is even more impressive when taking into account the nature of this program, since it is a RTO Task Group and not the primary job of the participants of the group. The interaction between predictions and experiments has already shown a great deal of value, since the predictions showed some of the problems with the initial design and greatly aided in the redesign of the configuration. Also the predictions aided in the wind tunnel test planning and provided estimates of loads for model design and areas of interesting aerodynamics for test scheduling. The comprehensive data being collected will prove invaluable to the researchers performing predictions, and will hopefully accomplish the initial goals of the group: to “understand the developing flow structures” and be able to accurately predict them for stability and control purposes. Initial design

Design of SACCON

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Configurations and wind tunnel models Within the AVT-161 Task Group two highly swept wing configurations are used for experimental and numerical investigations. The DLR X-31 wind tunnel model configuration is a double delta wing configuration with canard. The inner delta wing has a sweep angle of 57° and the outer 45°. The canard is a cropped delta wing with a sweep angle of 45°. Additional characteristics of the model are the inner and outer leading edge flaps, the trailing edge flaps, the front wing and rear fuselage strakes as well as the tail plane with rudder. The investigations within the AVT-161 task group are done without any control device deflection but taking the stream wise gaps between the leading edge flaps into account. Figure 6: X-31 low speed wind tunnel model on the The X-31 wind tunnel model configuration on the MPM-“Model Positioning Mechanism” in the open MPM system is shown in Figure 6. test section in the Low Speed Wind Tunnel The main target configuration of the task group is a Braunschweig (DNW-NWB). generic UCAV geometry called SACCON – “Stability And Control CONfiguration”. The plan form and section profiles are provided in cooperation with EADS-MAS. DLR adjusted the pre-design geometry for wind tunnel design purposes which actually lead into a higher overall thickness at the root chord to provide enough space for the internal strain gauge balance. Finally the wind tunnel model was designed and manufactured within the responsibility of NASA at NASA Langley Research Center. The model mounted on the MPM system in the low speed wind tunnel Braunschweig (DNW-NWB), is shown in Figure 7. The SACCON UCAV has a lambda wing plan form with a leading edge sweep angle of 53°, see Figures 8 and 9. The root chord is approximately 1m and the wing span is 1.53 m. The main sections of the model are the fuselage, the wing section and wing tip. The configuration is defined by three different profiles at the root section of the fuselage, two sections with the same profile at the inner wing, forming the transition from the fuselage to wing and the outer wing section. Due to radar signature issues the leading edge is parallel to the wing trailing edges and the wing tip is designed parallel to the trailing edge of the fuselage section. Finally the Figure 7: UCAV low speed wind tunnel model Saccon outer wing section profile is twisted by 5° around the leading edge to reduce the aerodynamic loads on the MPM-“Model Positioning Mechanism” in the and shifting the onset of flow separation to higher closed test section of the Low Speed Wind Tunnel angles of attack. Braunschweig (DNW-NWB). The wind tunnel model consists of different modular parts. The leading edge is exchangeable providing a sharp (SLE) and a variable round leading edge (RLE). The RLE is sharp at the root chord and the leading edge radius is growing in span wise direction up to the intersection between fuselage and wing and decreasing again. The overall radius distribution and relative thickness of the RLE configuration is given in Figure 10. Furthermore the wing tips can be separated and the rear part of the wing section is prepared to add control devices. The model is made of CFRP and is very light with an overall weight of less than 10kg (including pressure tubes and PSI modules). The very light design reduces the dynamic inertia loads enabling the use of a smaller, more sensitive balance that provides better force and moment resolution. The model was designed to accommodate a rear sting or belly sting mount for tests in the DNW-NWB low speed facility at DLR in Braunschweig and the 14′×22′ low speed wind tunnel at NASA LaRC. Different

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Figure 8: Plan form and geometric parameters of the UCAV Saccon configuration.

Figure 9: Wing sections and 3D view of the SACCON UCAV configuration.

connection links between belly sting support and internal balance at NWB provide an angle of attack range from -15° to 30°. This is provided by two different rigid cranked yaw links or by using an internal pitch link driven by a 7th axis. The two different basic setups with and without 7th axis are shown in Figure 11. The position of the belly sting connection is chosen to minimize the influence of the sting on the overall flow topology. This applies mostly to the vertical flow on the upper side of the model. This influence has to be approved via CFD simulations. Out of previous investigations with the X-31 configuration it was shown that for the prediction of the total forces and moments the sting support has to be taken into account [3].

Figure 10: Radius distribution and relative thickness of the RLE Saccon configuration. It can be seen in Fig. 6 that the connection between sting support and internal balance is completely covered by the model fuselage for the configurations with yaw link. These are adapted new designs especially for the SACCON configuration. For the pitch link it was not possible to adapt the design and a cover was used to smooth the geometry in this area. Figure 11: Top: 15° cranked yaw link support. The SACCON wind tunnel model is equipped with more Button: Support with 7th axis and internal pitch than 200 pressure taps on the upper and lower side of the link.

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model. The taps are connected with pressure tube to PSI modules within the model. At nine additional positions unsteady pressure sensors are mounted. The location of x/c = 0.2 the pressure taps is depicted in Figure 12. The pressure tap locations result from preliminary CFD computations LE prior to the wind tunnel tests. The aim was to capture the x/c = 0.45 ⊥ 6 0. = complex vortex flow topology over the configuration y/s = 0.26 y/s and operation points of the trajectory. One of these x/c = 0.7 LE simulations results are seen as well in Fig. 7. The ⊥ 94 0. preliminary simulations based on best practice = y/s procedures resulting from CFD prediction of configurations with round leading edges like results from AVT-113 [13]-[19]. All pressure tube connections Kulites between the pressure taps and PSI modules are of the Figure 12: Pressure tap location on the upper same length to guarantee the same time dependent surface of the Saccon configuration (Surface: behavior for each pressure taps during the unsteady pressure distribution, preliminary CFD pressure measurements. This leads to big bundles of the calculation). flexible tubing which have to be carefully installed to prevent kinks. The tubes bundles which have to be placed inside the model are shown in Figure 13. The finish of the model is motivated by PIV measurement requirements. The shiny black paint contains particles of Rhodamine B. By this the reflected light consists of a different wave length which can be filtered. This leads to a high accurate PIV measurement close to the surface [44]. Initial tests with the RLE configuration showed an arbitrary transition line on the upper surface of the model detected by infrared thermography. These measurements lead to the decision to prepare the leading edge with a carborundum grit trip as it is depicted in Figure 14. The difference between the configuration with and without transition tripping is shown in Figure 15. On the left hand side the clean configuration with no tripping at the round leading edge and an arbitrary transition line between laminar and turbulent flow can be observed. On the right hand side the infrared thermography picture after the installation of the carborundum grit trip is shown. It can be observed that after establishing the grit a fully turbulent flow over the upper Figure 13: Pressure tubes and PSI modules of the wing surface can be assumed. Saccon wind tunnel model. r

r

r

Arbitrary transition line SACCON RLE α = 5° β = 0° V = 60 m/s 0° yaw link

Figure 14: Leading edge with carborundum grit trip on the RLE-FT configuration (FT: fixed transition).

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Fully turbulent SACCON RLE-FT α = 5° β = 0° V = 50 m/s 15° Yaw link

Figure 15: Upper surface infrared thermography pictures of the Saccon configuration. Left: RLE with clean leading edge Right: RLE-FT with carborundum grit trip.


Experimental Approach The following chapter describes the wind tunnel facilities used for establishing the steady state and dynamic experimental data for the X-31 and SACCON configuration. The first part of the SACCON tests was conducted in the Low Speed Wind Tunnel Braunschweig (DNW-NWB) which is part of the foundation “German Dutch Wind Tunnels”. The second part of SACCON tests will be done at NASA Langley Research Center (LaRC) in the 14′×22′ low speed wind tunnel facility. Finally it is planned that at the Defense Science and Technology Organization (DSTO) in Australia some basic research will be performed focused on vertical flow structures and derivatives with a small scaled model in a water tunnel. Wind Tunnel Low Speed Wind Tunnel Braunschweig (DNW-NWB) The DNW-NWB belongs to the foundation “German–Dutch Wind Tunnels” under Dutch law. DNW operates 12 different wind tunnels on five sites in Germany and the Netherlands. The DNW-NWB is located on the DLR site in Braunschweig, Germany. It is a closed-circuit, atmospheric type wind tunnel, which can be operated either with an open, slotted or closed test section. A plan view of the tunnel is given in Figure 16. The test section size is 3.25 m by 2.8 m (10.6′ by 9.2′). The maximum free stream velocity is V=80 m/s (263 ft/s) in the closed test section and V=70 m/s (230 ft/s) in the open test section. High performance secondary air systems (compressors and vacuum pumps) allow for tests featuring turbine power simulation (TPS) or engine intake tests. NWB’s model supports include basic α−β−support, half-model support, support for 2-D-models, a rotary motion support for rolling and spinning tests and the Model Positioning Mechanism (MPM), which will be described in more detail in section B. 14′×22′ low speed wind tunnel The forced oscillation test rig in the NASA Langley 14-by-22-Foot Subsonic Tunnel (shown in Figure 17), can provide constant amplitude and frequency sinusoidal motion in the roll yaw or pitch axes [45]. The frequency can be set from 0.1 to 1.0Hz at amplitudes up to 30 degrees. The SACCON model is scheduled for roll and yaw oscillation tests in this tunnel in the fall of 2009. The yaw oscillation testing will repeat some of the conditions tested in the DNW-NWB Low Speed Tunnel for test to test comparison.

Figure 16: Low speed wind tunnel DNW-NWB Braunschweig.

Figure 17: Model setup in the 14′×22′ low speed wind tunnel at NASA LaRC.

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The Model Positioning Mechanism

Figure 18: Saccon on the MPM support in the DNW-NWB wind tunnel.

DNW-NWB’s Model Positioning Mechanism is a six DOF parallel kinematics system designed for static as well as for dynamic model support. Characteristic features of this unique test rig are the six constant length struts of ultra high modulus carbon fibre and the six electric linear motors, which move along two parallel rails. The first Eigenfrequency at the MRP is above 20 Hz. The MPM is located above the test section and can be operated in the open test section as well as in the closed one. The location of oscillation axes can be chosen arbitrarily and in addition to classic sinusoidal oscillations the MPM can perform multi-DOF maneuvers. An artist’s impression of the MPM carrying the SACCON model in the open test section is given in Figure 18. More details concerning the MPM are given in Bergmann et al. [46][47].

Measurement technique Transition Because the precise knowledge of the flow conditions is of paramount importance for tests, which are used for CFD validation, the wind tunnel entry was begun with boundary layer transition observations using infra red thermography. In case the model surface and the passing air have different temperatures, the transition line can be observed with an infra red camera of suitable sensitivity because of the different heat transfer properties of laminar and turbulent boundary layers. The DLR Institute of Aerodynamics and Flow Technology utilizes a set of cameras on remotely controllable mounts for this purpose since more than a decade. Typical results are given in Fig. 10. More information are given by de Groot [48] and Fey et al. [49]. Forces and Moments For the measurements of forces and moments internal six component strain gauge balances have been used. The X-31 wind tunnel models have been equipped with an Emmen 192-6 balance, whereas the SACCON model has been instrumented with a smaller Emmen 196-6 balance. Pressures The X-31 wind tunnel models are equipped with 54 in situ mounted Kulite pressure transducers. To permit pressure measurements on DNW-NWB’s rolling and spinning support the transducers are of the absolute pressure measurement type. The Kulites are not installed permanently but can be dismounted, so one set of transducers can be used for both X-31 wind tunnel models. For the SACCON wind tunnel model a combination of in situ mounted Kulite pressure transducers and 48channel ESP modules, mounted in the model’s center section have been used. The Kulite transducers have been installed to verify the synchronicity of the ESP modules with respect to balance and position data acquisition. The signals of the ESP modules have been corrected for attenuation and phase shift according to Nyland et al [50]. Model Position For the measurement of the instantaneous attitude and position of both X-31 and SACCON wind tunnel models DNW-NWB employed an optical system featuring two high speed video cameras. The cameras have been mounted below the test section and acquire SXGA (1280x1024) images at 300 frames per second, each. The position of the model is calculated in real time from the pixel co-ordinates of three markers, which have been applied to the model surface. For static angle of attack measurement a conventional inclinometer has been installed additionally. Two different data acquisition systems have been employed for force and pressure measurements in the X-31 tests on the one hand and the SACCON tests on the other hand. The X-31 test has been performed with a

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telemetric system consisting of 64 channels with a 16 bit resolution. The SACCON test has been made with DNW-NWB’s standard Hottinger-Baldwin MGC+ data acquisition. PIV- “Particle Image Velocimetry” Three component PIV measurements have been performed on the SACCON wind tunnel model in the DNW-NWB by teams of ONERA and DLR [51] simultaneously. Cameras and laser light sheets have been installed outside the test section without direct mechanical contact to the test section structure in order to avoid quality reduction due to vibration. Optical access was granted by windows in the test section side walls, as can be seen in Figure 19. Cameras and laser light sheets were aligned parallel and orthogonal to the model’s x-axis at an angle of attack of α = 17°, as the force and pressure measurements indicated that an angle of attack range between 14° and 20° offers the most interesting and challenging flow topologies for CFD. The Figure 19: PIV measurements of the Saccon ONERA PIV setup was oriented towards the front part configuration in the closed test section at whereas the DLR PIV setup was oriented toward the model’s DNW-NWB Braunschweig. rear part. Employing the MPM’s capabilities, the model was displaced along its x-axis in order to move the light sheets along the model. This eliminated the need for time consuming re-calibrations of the cameras when the location of the light sheet relative to the model was changed. While most of the PIV measurements have been made on the right hand side of the model, the DLR team also made measurements on the left hand side. Figure 20 illustrates the locations of the PIV light sheets on the SACCON model.

Figure 21: Example of PIV results in selected planes obtained by DLR giving insights in the flow topology above the UCAV model.

Figure 20: Positions and AoA for PIV measurements.

The DLR PIV team used 300 images from each camera. The ONERA team used 800 images. The ONERA cameras had a higher resolution than the DLR cameras but the oberserved area was smaller. A typical PIV result is given in Figure 21. In addition to the static PIV measurements also measurements on the model oscillating in pitch have been performed. The PIV cameras were triggered in a phase locked mode by the model oscillation. PIV data are taken at eight different phase angle over the oscillation cycle. For each phase angle fifty measurements ate taken averaging.

12


Wind tunnel tests Within AVT-161 two additional wind tunnel campaigns with the X-31 configuration in the open test section of the DNW-NWB were done. The first test focused at the influence of the model setup on the integral data and pressure distribution on the model. For instance the influences of the sting support, of gaps between the support and the model as well as on flap gaps and elastic effects at the hinge line of the leading and trailing edge flaps. In Figure 22 the different setups related to the leading edge flap gaps are depicted. Three different setups where tested: All gaps sealed, all gaps open and only the span wise gaps sealed between the LE-flaps and the wing. The results for five different angles of attack are shown in Figure 23. It can be observed that there are significant differences in the pressure distribution for all three setups. This is caused by the fact that the vortex V = 60 m/s, no WT wall corrections, ηLEflaps = 0°

Influence of the LE flap gaps α = -5° α = 0° α = 5° α = 10° α = 15°

cp

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Figure 23: X-31: Influence of the LE flap gaps on the pressure distribution at five different angles of attack. formation is very sensitive to the setup at the LE. This effect has a minor influence on the overall lift but has significant influence on the longitudinal and lateral stability. The influence of the belly sting on the pitching moment is shown in Figure 24. In Fig. 24 the pitching moment verses the AoA is plotted. It can be observed that the additional 7th axis behind the belly

X31 R/C model w/o wall corrections

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Figure 22: X-31 wind tunnel model. Top: Position of the gaps at the Canard and control devices. Bottom: Picture of the gaps between the two LE flaps and between the flaps and the wing (view from below).

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Figure 25: X-31: Example for pitching maneuvers.

sting causes a parallel shift to higher values which causes a rear loading in comparison to the setup without 7th axis. A comparable setup can be seen in Fig. 11 for the SACCON configuration. The result from the first X-31 test led to the final configuration setup of the model and support for the second wind tunnel entry. In the second wind tunnel entry steady state and unsteady experimental data have been

13


obtained. The dynamic tests include sinusoidal pitch and yaw motion. These data were finally used for computer code validation purposes. First results of these investigations are described by Boelens [52], Jirasek and Cummings [53]. Two examples of the maneuvers which have been performed are shown in Figure 25. The red curve shows the AoA versus time of an overlay pitching maneuver. The MPM establishes the slower movement with high amplitude whereas the 7th axis is used to overlay the higher frequency motion with smaller amplitude. With the SACCON configuration two wind tunnel tests in the closed test section of the DNW-NWB are done up to now. A third test will follow at the end of 2009 in the 14′×22′ low speed wind tunnel at NASA LaRC to complete the experimental investigations. The first tests focused on the identification of the basic aerodynamic behavior of the SACCON configuration. The tests were done with three different belly sting setups. Due to the experiences of the X-31 tests two different belly sting supports are established to reach higher angles of attack without the influence of a 7th axis behind the sting. Nevertheless there is an influence of the support left just by changing the crank angle between the sting and the model by 15°, as seen in Figure 26. Fig. 26 shows as well the different AoA ranges which can be achieved by using the one or the other yaw link setup. Figure 27 shows the influence using the pitch link with 7th axis instead of the 15° cranked yaw link, see as well Fig. 11. Whereas SACCON 0° yaw link SACCON 15° yaw link

SACCON 15° yaw link SACCON 6° pitch link

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Figure 26: Saccon: Lift and pitching moment coefficient versus AoA. Influence of the yaw link crank angle.

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Figure 27: Saccon: Lift and pitching moment coefficient versus AoA. Difference between 15° cranked yaw link and internal pitch link.

in Fig. 24 and Fig. 26 only a parallel shift of the pitching moment coefficient can be observed due to a change of the setup, is the influence of the 7th axis setup in Fig. 27 is different. This different behavior is probably caused by the additional cover between the sting and the model fuselage. Figure 28 and 29 should show the aerodynamic characteristic behavior of the SACCON configuration itself. In Fig. 23 the lift and pitching moment coefficient versus AoA for sharp LE, round LE and round LE with fixed transition is shown. It can be seen that the overall lift is fairly the same for all kinds of leading edges. Only the curve of the sharp leading edge configuration shows a higher lift at AoA above 12°. This is caused by the higher nonlinear component of lift due to stronger vortices which occur at the sharp leading edge. The pitching moment coefficient shows much more sensitivity due to the change of the LE shape. The characteristic discontinuous behavior between AoA of 10° and 20° takes place earlier for the sharp leading edge as for the two round leading edge configurations. RLE 0° yaw link V = 60 m/s SLE 0° yaw link V = 60 m/s RLE-FT 0° yaw link V = 50 m/s

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Figure 28: Saccon: Lift and pitching moment coefficient verses AoA for different leading edges.

Cmz

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Figure 29: Saccon: Side force, roll moment and yaw moment coefficient verses AoA for different angles of side slip.

14


Fig. 28 shows as well a significant different pitching moment distribution between the RLE and RLE-FT configuration. The reason for this behavior is to some extent already explained by the task group taking experimental data and CFD results into account. The discussion of these results should not be part of this publication and will be delivered later. Fig. 29 shows the behavior of the SACCON configuration at nonsymmetric on flow conditions. It is seen that at AoA higher than 10° depending on the direction of the angle of sideslip significant differences can be observed. Even at small angles of side slip this non-symmetric behavior can be seen. To ascertain the flow causing these flow conditions is part of the task group and one major issue of the code validation process to get the ability to predict seriously the stability and control issues of the target configurations. All the discussed parameters above have to be taken into account when preparing the computational simulations. Further boundary conditions are, among others, the condition of the model surface (taped screws and gaps, surface roughness), transition effects, as discussed before in Chapter III, and the wind tunnel corrections. In the 14′×22′ low speed wind tunnel at NASA LaRC it is proposed to complete the test matrix related to SACCON configuration. This includes some replication of the static and yaw oscillation runs for tunnel to tunnel comparisons followed by a series of roll oscillation runs at different frequencies and amplitudes. As with the DNW-NWB tests the time history data of the pressures as well as the forces and moments will be recorded.

Dynamic Data Analysis As was previously noted the dynamic data runs consisted of a 30 second record of the measured forces, moments, model position and pressures. Each multi-cycle, sinusoidal data run was later condensed to a one cycle average with standard deviation. Figure 30 shows an example of a 1Hz pitch oscillation pitching moment data set for the SACCON RLE-FT configuration. The nominal angle of attack is 20° with a pitch amplitude of Δα=5°. Also shown in the figure are the static data and the 1st-cycle average with standard deviation bars. Nominally 5 data runs were taken at each test condition. Figure 30: Saccon example of pitch oscillation The pitch dynamic effects on the lift and pitching moment 1-cycle average pitching moment coefficient of the SACCON RLE-FT configuration are shown in verses AoA. Figures 31 and 32 for 1 and 3Hz oscillation frequencies, respectively. The 5° amplitude oscillations are shown about nominal angle of attacks of 5°, 10°, 15° and 20°. The dynamic damping effect is seen as the difference between the static and dynamic measurements. A linear damping effect with sinusoidal oscillations results in an ellipsoid about the static data which grows with frequency, as seen in Fig. 32. The pitching moment showed a greater dynamic effect than the lift which showed very small effect below 15° AoA. In the higher AoA range near flow separation the lower frequency (1Hz) resulted in a more non-linear behavior than the higher 3 Hz data. This is presumed to be due to the flow dynamics having sufficient time to transition between states at the lower frequencies. At the higher frequencies the flow does not have time to transition Figure 31: Saccon RLE-FT Figure 32: Saccon RLE-FT 1Hz pitch oscillation lift and 3Hz pitch oscillation lift and resulting in a more linear behavior. The greater CFD challenge is in capturing the pitching moment dynamic pitching moment dynamic lower frequency non-linear dynamic data about 5°, 10°, 15° and data about 5°, 10°, 15° and 20° nominal AoA.

20° nominal AoA.

15


effect. The effect of the SLE on the pitch dynamics is shown in Figure 33 for the 1Hz pitch oscillation runs. The dynamic effects are similar to the RLE-FT results shown in Fig. 31. An example of the yaw oscillation dynamic effects on the rolling and yawing moment coefficients of the SACCON SLE configuration about a 15° nominal AoA are shown in Figure 34. The lack of any vertical surfaces on the SACCON configuration results in fairly small yaw oscillation effects. The asymmetry of the data at this AoA is indicative of small asymmetries in the model geometry which was also evident in Fig. 29.

Figure 33: Saccon SLE 1Hz pitch oscillation lift and pitching moment dynamic data about 15°, 20° and 25° nominal AoA.

Figure 34: Saccon SLE yaw oscillation rolling and yawing moment dynamic data about 15° nominal AoA.

Numerical Approach The numerical simulations will be done by several partners using various kinds of RANS solving computational codes using cell-centered or cell-vertex metric, fully structured or hybrid grid approaches. Additionally to that some contributions are given using engineering methods for RANS solver comparisons in areas of linear aerodynamic behavior. Furthermore different “best practices” experiences exist for each computational method which leads to the necessity to define common procedures or courses of action to be able to compare the different approaches and being able to detect similarities and differences. Guidelines need to be defined, similarities and differences between various approaches have to be detected to develop advanced “best practice” procedures or to identify possible improvements of particular “best practice” procedures. In AVT-161 common test cases were defined based on pre-design CFD calculations which detected the AoA range of interest. Then the participants confirmed to perform CFD simulations using two different turbulent model approaches (Spalart-Allmaras and Wilcox k-ω) which should give the ability for cross comparisons even

Figure 35: X-31: CFD NLR EnSolv solution. Pressure distribution and vortical flow structure at an AoA of α = 20° (Low speed wind tunnel conditions) by Boelens [52].

Figure 36: Saccon: CFD DLR TAU-Code solution. Pressure distribution and vortical flow structure at an AoA of α = 15° and bank angle of β = 5° (Low speed wind tunnel conditions).

16


though the methods are more or less different in there overall implementation. Although the basic equations are similar for the turbulence models, the implemented boundary conditions should be given for each method to identify differences related to this issue. Finally the influence of the computational grid between each computational approach should be identified by analyzing the grid characteristics (Spread sheet of: Initial spacing, resolution in stream wise and span wise direction, resolution of the boundary layer, resolution of the near field of the model, etc.). Due to the huge amount of experimental data and with respect to time and resources consuming unsteady simulations common dynamic test cases are defined. Even the order of the common test case simulations were determined to enable as much as possible cross comparisons at a time. Figure 35 and 36 show two preliminary CFD results from the X-31 [52] and the SACCON configuration. The two examples show the major impact of the vertical flow on the flow topology on the upper side of the configurations.

Conclusions A NATO RTO Task Group, AVT-161, is utilizing state-of-the-art computational and experimental tools to determine the ability to accurately predict the aerodynamics of maneuvering aircraft and ships for stability and control purposes. This is being accomplished for aircraft by investigating the static and dynamic aerodynamics of two vehicles: the X-31 and SACCON, a generic UCAV. The integrated approach for conducting this research has been described, including the initial planning of the wind tunnel tests and the design of SACCON configuration, which incorporated CFD, lower-order prediction methods, and wind tunnel capabilities. This ongoing research program has determined the importance of knowing and defining conditions in the tunnels, since this provides the boundary conditions for the numerical predictions. This includes the freestream conditions and any and all corrections that are made to the data. The CFD researchers have been actively involved in the planning of the test, and have all of the details of the experiments readily available via a Task Group website. In addition, common test cases (both static and dynamic) have been identified for both aircraft to form a basis for comparing the various computational and modeling approaches. This approach has shown itself to greatly enhance the interaction among the various sub-teams of the Task Group as we predict, measure, and analyze the flight mechanics of these aircraft. Detailed results for the aircraft will be compared between researchers, including both numerical predictions and experimental data, on common cross plots so that discussions and knowledge building can take place among the team members. While our overall goal is to determine the state-of-the-art for computational capabilities in predicting stability and control parameters for aircraft, we are also conducting detailed “experiments” to assess the ability of grids, turbulence models, and time integration approaches to accurately predict these complex unsteady flow fields. The end product of this work should greatly increase our understanding of both aerodynamics and our ability to predict complex flow fields with various approaches.

Acknowledgments The authors would like to thank NASA Langley and Advanced Technologies Incorporated for designing, manufacturing and providing the SACCON wind tunnel model to the AVT-161 task group as well as DLR for providing wind tunnel test time within the DNW-NWB for SACCON and X-31 tests. Also the authors would like to thank the colleagues from ONERA Lille and from DLR Göttingen for the PIV investigations, especially Anne Gilliot, Andreas Schröder and Robert Konrath. Furthermore we would like to thank Thomas Weddig, Jan Himisch and Carsten Liersch from DLR for their investigations on the SACCON configuration within the predesign phase. Marvin Gülzow, Martin Sitzmann and Claus-Peter Krückeberg from DLR for preparing the wind tunnel model for testing. Klaus de Groot and Saskia Hempel for transition measurement support. Finally, thanks for all the contributions by the highly motivated participants of AVT-161 task group.

17


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