Ünver Kaynak

Advances in the computation of transonic separated flows over finite wings

Abstract Transonic flow fields around a low-aspect-ratio wing in a wind tunnel have been simulated using an Euler/Navier-Stokes zonal approach. By using a fast parabolic-type grid generator, a global grid with an H-H topolgy is generated around the wing in which the far-field boundaries match the tests section of the wind tunnel. This global grid, which is a coarse grid, was then subdivided into separate zones. The grid zones near the wing were suitably clustered for viscous resolution, and the Reynolds-averaged Navier-Stokes equations were solved. In the rest of the flow field the Euler equations were solved. Particular emphasis was placed on simulating shock-induced separated flows. Problem areas in the state-of-art computational techniques in wind-tunnel simulations were identified, and each problem was studied in an isolated manner in order to avoid combined effects of multiple sources. The lessons that were learned from those studies were then combined to yield more accurate wind-tunner simulations and to further the understanding of separated flows. As a technology demonstration, one solution with extremely fine grid resolution (1.1 × 10 6 total grid points) was achieved using the Cray-2-superdomputer. Of particular importance for flow physics, a transonic mushroom-type separated flow with counterrotating vortices that closely resembles the experimental pattern, was demonstrated for the first time.

Design of a Flight Stabilizer System for a Small Fixed Wing Unmanned Aerial Vehicle using System Identification

Abstract Flight stabilizers for unmanned aerial vehicles (UAVs) provide level flight and make the UAVs more reliable and operational. In this paper a system identification process is applied to flight data logs obtained from flight simulations and a stabilizer system is designed using the dynamical model obtained from system identification. Hardware-in-the-loop testing was also performed on the flight stabilizer system designed is also tested using a custom built moving platform in the pitch and roll axis with actual flight sensors.

Hardware-in-the-Loop Test Platform Design for UAV Applications

A hardware-in-the-loop (HIL) platform for unmanned air vehicle (UAV) systems is designed that demonstrates flight attitudes on yaw, pitch and roll axes. The design combines a sophisticated flight simulation software with a platform capable of moving 360 degrees on all axes. This enables the testing of the flight sensors and autopilot algorithms for all sorts of scenarios including emergency and acrobatic cases where an indefinite number of full rotations in the yaw, roll and pitch might take place.

Automatic Recovery and Autonomous Navigation of Disabled Aircraft After Control Surface Actuator Jam

Loss of flight control authority in result of a system damage or an actuator jam is known or suspected to be the potential cause of many aviation accidents. After losing one or several control surface actuators, sustaining the flight for a safe return home depends on the capability of the remaining surfaces’ control authority and the engine power. In this study, automatic recovery and autonomous guidance of a disabled general aviation aircraft is demonstrated. A nonlinear aircraft model and a MATLAB/SIMULINK based flight dynamics and control toolboxes are used to develop flight control laws for the disabled aircraft. The flight control laws are first validated for steady trimmed flight conditions and an actuator jam is applied to demonstrate the automatic recovery and autonomous navigation of the aircraft. The autopilot is designed to handle different control actuator malfunctions including rudder or aileron jams. Different scenario based simulations show that the new autopilot design is capable of sustaining safe flight and autonomous navigation under such malfunctions.

Design of a flight stabilizer for fixed-wing aircrafts using H∞ loop shaping method

Autopilot systems for unmanned aerial vehicles (UAVs) and aircrafts provide flight missions without need of human input and make them more reliable and efficient. The first step of designing an autopilot is a stabilizer mode. Conventional autopilot systems have inner and outer loops. Stabilizer is the inner loop for an autopilot. In this paper designing an aircraft control system ensures good performance and robustness that allows control of roll, pitch and yaw angles will be declared. Aircraft dynamics are used to design the model of the control system in the MATLAB Simulink environment. Using loop shaping method to achieve stable and robust control system will be the strategy. Generated controllers to find the most effective one are embedded in the X-plane which is one of the realistic simulations.

HARDWARE-IN-THE-LOOP TEST PLATFORM FOR A SMALL FIXED WING UNMANNED AERIAL VEHICLE EMBEDDED CONTROLLER

Hardware-in-the-loop (HIL) test systems for unmanned aerial vehicles (UAV) embedded flight controllers provides more realistic test environments for pre-flight tests. The HIL test system which is described in this paper combines a detailed and sophisticated flight simulation software with a moving hardware platform in the pitch and roll axis with actual flight sensors to form a complete testing environment.

A Novel Intermittency Distribution Based Transition Model For Low-Re Number Airfoils

A new correlation based transition model is proposed using a novel intermittency distribution function that relies on local information. The intermittency behavior of the transitional flows is reflected into computations by multiplying the production term of the Spalart-Allmaras turbulence model with the new intermittency factor. In this way, the model is in a sense similar to an algebraic model by using only an intermittency function rather than an intermittency transport equation, yet it carries transport equation character by means of the one-equation turbulence transport equation. Therefore, the present formulation being a local model yet bringing in the correlation data achieves a similar effect by using two less equations than similar transport equation based transition models like Menter’s model. Validation of the new model with known flat plate test data of Schubauer & Klebanoff and Savill shows quite good agreement with the experiment. Model was also tested against some moderately high Reynolds number airfoil cases and some low Reynolds number airfoil cases with very promising results. The results imply that the new model may become a viable alternative for the higher order methods that is especially attractive in the design environment.

Transition at Low-Re Numbers for some Airfoils at High Subsonic Mach Numbers

‡High altitude long endurance UAV flight regime imposes certain difficult testing conditions for the ground based test facilities, such as low density, low freestream turbulence, high alpha, low-Reynolds, and highsubsonic-Mach numbers. High altitude flight testing would be required to collect actual experimental data at an expense of much higher costs. Computational approach is a viable alternative to support generation of design data base for the laminar and transitional boundary layers at high-subsonic-Mach-numbers. In the present study, a two-equation γ-Reθt correlation-based transition model is further developed to predict some airfoils that are frequently used in the design of UAVs. Firstly, the empirical correlations already validated for low to high Mach number flat plate cases is validated for the thin NACA64A006 airfoil at low subsonic speed and high-Reynolds-number. Secondly, the present methodology is successfully demonstrated for the E387 and SD7037 moderately thick UAV type airfoils in the low-Reynolds and low-subsonic-Mach-number conditions. Finally, the relatively thicker APEX-16 airfoils at high-altitude, low-Reynolds-number conditions for high-subsonic Mach numbers are simulated. Results are compared with the available numerical data in the literature obtained through Reynolds Averaged Navier-Stokes and viscous-inviscid interaction methods using the e N method.

Control of Flow Separation and Transition Point over an Aerofoil at Low Re Number using Simultaneous Blowing and Suction

Control of flow over a NACA2415 aerofoil which experiences a laminar separation bubble for a transitional Reynolds number of 2x10 is numerically simulated under the effects of blowing and suction. An earlier experimental study using hot-wire anemometry for a clean (no jet) NACA 2415 aerofoil at α = 8° shows a laminar separation bubble over the one-third of the airfoil upper surface. In the no jet case, the recently developed k-kL-ω and k-ω SST transition models accurately predict the location and extent of the separation bubble. Later, single or multiple jets with a width of 2.5% the chord length are placed on the aerofoil upper surface for simulating the isolated or simultaneous blowing and suction jets. For single jets with blowing or suction, whereas the k-ω SST model suppress the separation bubble in both cases, the k-kL-ω transition model doesn’t completely eliminate the separation bubble but moves it downstream in the suction case. Overall, the blowing/suction control mechanism appears to be suppression of the separation bubble and reduction of the upper surface pressure to increase the lift and decrease the drag. For simultaneous blowing and suction, firstly, a blowing jet at 10%c and a suction jet at 36%c are placed at the beginning and end of the separation bubble based on the best results of the single jets. Then, the locations of the blowing and suction jets are reversed. In both cases, both transition models eliminate the separation bubble resulting in increase in the lift and decrease in the drag. However, the best results in terms of the L/D ratios are still obtained with the single suction jets.

A correlation-based algebraic transition model:

A correlation-based algebraic transition model that relies on local flow information is proposed. The model is qualified as an algebraic model, or a zero-equation model since it includes an intermittency function in place of an intermittency equation that is found in one- or two-equation models. The basic idea behind the model is that, instead of deriving new equations for intermittency transport, existing transport terms of the Spalart–Allmaras (S-A) turbulence model can be used. To this end, the production term of the S-A model is multiplied with the proposed intermittency function γBC; thereby the turbulence production is damped until it satisfies some turbulence onset requirements. The proposed formulation also depends on local information that uses empirical correlations to detect the transition onset using less equations and less calibration constants than other higher order models. The model is first validated against some widely-used zero and variable pressure gradient flat plate test cases with quite successful results. Second, the model is employed for some low Reynolds number airfoil cases with very promising results. Third, the model is applied for a turbine cascade case with success. Finally, two different three-dimensional wing flow cases were calculated under transonic and low subsonic flow conditions. To this end, the DLR-F5 wing subject to a transonic Mach number of 0.82 and the low-speed NREL wind turbine flow case are simulated and good agreement with experiments are observed. The results indicate that the proposed model may become an alternative for other models as it uses less computational resources with equivalent or higher accuracy characteristics that is quite advantageous for the computational fluid dynamics design in industry.