Okko J. Boelens

Lessons Learned from Numerical Simulations of the F-16XL Aircraft at Flight Conditions

This thesis covers the field of vortex-flow dominated external aerodynamics. As part of the contribution to the AVT-113 task group it was possible to prove the feasibility of high Reynolds number CFD computations to resolve and thus better understand the peculiar dual vortex system encountered on the VFE-2 blunt leading edge delta wing at low to moderate incidences. Initial investigations into this phenomenon seemed to undermine the hypothesis, that the formation of the inner vortex system depends on the laminar/turbulent state of the boundary layer at separation onset. As a result of this research, the initial hypothesis had to be expanded to account also for high Reynolds number cases, where a laminar boundary layer at separation onset can be excluded. In addition, unsteady transonic computations are used to shed light on a highly non-linear phenomenon encountered at high angles of incidence. At certain conditions, the increase of the incidence by a single degree leads to a sudden movement of the vortex breakdown location from the trailing edge to mid-chord. The lessons learned from the contribution to the VFE-2 facet are furthermore used to prove the technology readiness level of the tools within the second facet of AVT-113, the Cranked Arrow Wing Aerodynamics Project International (CAWAPI). The platform for this investigation, the F-16XL aircraft, experiences at high transonic speeds and low incidence a complex interaction between the leading edge vortex and a strong, mid-chord shock wave. A synergetic effect of VFE-2 with a further project, the Environmentally friendly High Speed Aircraft (HISAC), is also presented in this thesis. Reynolds number dependence is documented in respect to leading edge vortex formation of the wing planform for a reference HISAC configuration. Furthermore, proof is found for a similar dual vortex system as for the VFE-2 blunt leading edge configuration.

Boundary Conforming Discontinuous Galerkin Finite Element Approach for Rotorcraft Simulations

A numerical method has been developed for predicting the complex vortex wake for a helicopter rotor in hover and in forward flight. The method is based on the solution of the three-dimensional, compressible Euler equations expressed in an arbitrary Lagrangian Eulerian reference frame. A second-order accurate discontinuous Galerkin finite element method is used to discretize the governing equations on a hexahedral mesh. Unstructured, local mesh refinement is performed to enable prediction of the structure of the vortex wake. The capabilities of this computational fluid dynamics method are demonstrated by simulations of the flow around the Caradonna-Tung helicopter rotor in hover and simulations of the flow around the Operational Loads Survey helicopter rotor in forward flight

Shock Effects on Delta Wing Vortex Breakdown

It has been observed that delta wings placed in a transonic freestream can experience a sudden movement of the vortex breakdown location as the angle of incidence is increased. The current paper uses computational fluid dynamics to examine this behavior in detail. The study shows that a shock/vortex interaction is responsible. The balance of the vortex strength and axial flow and the shock strength are examined to provide an explanation of the sensitivity of the breakdown location. Limited experimental data are available to supplement the computational fluid dynamics results in certain key respects, and the ideal synergy between computational fluid dynamics and experiments for this problem is considered.

Vortical flow prediction validation for an unmanned combat air vehicle model

As part of the NATO Applied Vehicle Technology 161 technical group, a study of the aerodynamic behavior of the stability and control configuration wind-tunnel model is presented. Sharp and round leading-edge versions of the model are computed. A validation of Reynolds-averaged Navier–Stokes predictions obtained using two block structured codes are made. Static cases are analyzed and compared with wind-tunnel measurements. The vortical flow features are described in detail for a range of angles of attack. The predictions are in good agreement with the experiments at low angles of attack, whereas for higher angles of incidence (alpha > 15), some discrepancies are seen. A dual vortex structure is present in this region for both leading-edge configurations, resulting in highly nonlinear aerodynamic behavior.

Analysis of Transonic Flow on a Slender Delta Wing Using CFD

The behaviour of the flow over slender delta wings under transonic conditions is highly complex. With the occurrence of a number of shocks in the flow the behaviour of vortex breakdown is quite different to that for subsonic flow. This investigation considers this behaviour over the 65 sharp leading edge delta wing used in the 2nd International Vortex Flow Experiment (VFE-2) using Computational Fluid Dynamics. Three institution involved in the VFE-2 have collaborated to consider the wing under conditions of M = 0.85 and Re = 6 × 10 at two incidences: α = 18.5 and 23. The flow solutions are compared to existing experimental data and show good agreement for the cases considered. However, a discrepancy with the experimental data is shown where the critical incidence for the onset of vortex breakdown on the wing is under-predicted. From analysis of the solutions, it is determined that the onset of vortex breakdown is highly dependent on the vortex strength and the strength and location of the shocks in the flow. The occurrence of a critical relationship between these parameters is suggested for vortex breakdown to occur and is used to explain the discrepancies between the computational and experimental results based on the under-prediction of the vortex core axial velocity. A sensitivity study of the flow to a number of computational factors, such as turbulence model, is also undertaken. However, it is found that these parameters have little effect on the overall behaviour of the transonic flow.

Static and dynamic numerical simulations of a generic UCAV configuration with and without control devices

A contribution for the assessment of the static and dynamic aerodynamic behavior of a generic UCAV configuration with control devices using CFD methods is given. For the CFD simulations the unstructured grid based DLR TAU-Code and the structured grid based NLR solver ENSOLV are used. The numerical methods are verified by experimental wind tunnel data. The current investigations should provide a contribution to assess the prediction capability of control device effectiveness using CFD methods. The presented computational results for the assessment will be validated by dedicated experimental data. Furthermore, it should support the understanding of the flow physics around the trailing edge control devices of highly swept configurations with a vortex dominated flow field. Design requirements should be able draw by analyzing the interaction between the vortical flow and the control devices. The present work is part of the NATO STO/AVT Task Group AVT-201 on Stability and Control prediction methods

Comparison of measured and simulated flow features for the full-scale F-16XL aircraft

Accurate and cost-effective Computational Fluid Dynamics (CFD) methods play an increasingly important role, even in the support of fighter aircraft operations. Prior to the deployment of such CFD methods they should be well validated and evaluated against stateof-the-art wind tunnel and/or flight test data. The Cranked-Arrow Wing Aerodynamics Project (CAWAP) provided the CFD community with an excellent database for validation and evaluation. Initiated by NASA, the Cranked-Arrow Wing Aerodynamics Project International (CAWAPI) was started as a follow-on project of CAWAP. The National Aerospace Laboratory NLR participated in this project using the in-house developed flow simulation system ENFLOW, which includes both grid generation tools and a flow solver. NLR applied (semi-automatic) grid generation tools to generate a structured (multi-block) grid. Steady flow simulations for all seven CAWAPI flight conditions are performed employing the flow solver ENSOLV. Results obtained for flight condition 7, 19 and 25 are discussed. The focus of this discussion is on a com parison of the measured and simulated flow features. It is shown that the understanding o f NLR’s structured (multi-block) grid generation algorithm and the confidence in the appl ication of its flow simulation method to complex fighter configurations increased significan tly by participating in CAWAPI.

What was learned from numerical simulations of F-16XL (CAWAPI) at flight conditions

Nine groups participating in the CAWAPI project have contributed steady and unsteady viscous simulations of a full-scale, semi-span model of the F-16XL aircraft. Three different categories of flight Reynolds/Mach number combinations were computed and compared with flight-test measurements for the purpose of code validation and improved understanding of the flight physics. Steady-state simulations are done with several turbulence models of different complexity with no topology information required and which overcome Boussinesq-assumption problems in vortical flows. Detached-eddy simulation (DES) and its successor delayed detached-eddy simulation (DDES) have been used to compute the time accurate flow development. Common structured and unstructured grids as well as individually-adapted unstructured grids were used. Although discrepancies are observed in the comparisons, overall reasonable agreement is demonstrated for surface pressure distribution, local skin friction and boundary velocity profiles at subsonic speeds. The physical modeling, be it steady or unsteady flow, and the grid resolution both contribute to the discrepancies observed in the comparisons with flight data, but at this time it cannot be determined how much each part contributes to the whole. Overall it can be said that the technology readiness of CFD-simulation technology for the study of vehicle performance has matured since 2001 such that it can be used today with a reasonable level of confidence for complex configurations.

Numerical and Theoretical Considerations for the Design of the AVT-183 Diamond-Wing Experimental Investigations

A diamond-wing configuration has been developed to isolate and study blunt-leadingedge vortex separation with both computations and experiments. The wing has been designed so that the results are relevant to a more complex Uninhabited Combat Air Vehicle concept known as SACCON. The numerical and theoretical development process for this diamond wing is presented, including a view toward planned wind tunnel experiments. This work was conducted under the NATO Science and Technology Organization, Applied Vehicle Technology panel. All information is in the public domain.

Flow Analysis of the F-16XL Aircraft (CAWAPI-2) At Transonic Flow Conditions

In the framework of the Cranked-Arrow Wing Aerodynamics Project International 2 (CAWAPI-2), a cooperation between Cassidian/EADS (Germany), KTH (Sweden), Lockheed-Martin (USA), NASA Langley (USA) and National Aerospace Laboratory NLR (The Netherlands) to further assess Computational Fluid Dynamics codes against F-16XL flight test data, National Aerospace Laboratory NLR performed an analysis of a set of transonic flight conditions available in the CAWAP database. Flight condition FC70 was used previously during the Cranked-Arrow Wing Aerodynamics Project International to investigate transonic flow on the F-16XL aircraft. During this project it was discovered that flight condition FC70 was flown with a deflected leading edge flap. To facilitate CFD analysis a transonic flight condition without deflected control surfaces was judged desirable by the CAWAPI-2 members. Therefore, it was decided to search the CAWAP database for a transonic flight condition without any control surface deflections. Since no information on the control surface deflections of the other transonic flight conditions was available to the CAWAPI-2 partners, an alternative approach to scrutinize the available flight test sectional surface pressure measurements for indications of control surface deflections was adopted. This analysis revealed that flight condition FC79 was the most likely transonic flight condition to be flown without any control surface deflections. Flight conditions FC70 and FC79 were analyzed using NLR’s in-house developed flow simulation system ENFLOW. Comparison of the measured and simulated surface pressure coefficients confirmed that flight condition FC79 was flown without any control surface deflections, and that this flight condition thus is much better suited for further comparisons between flight test data and CFD simulations.