Carl P. Tilmann

Design and application of compliant mechanisms for morphing aircraft structures

Morphing aircraft structures can significantly enhance air vehicle performance. This paper highlights ongoing work to design novel compliant mechanisms that efficiently morph aircraft structures in order to exploit aerodynamic benefits. Computational tools are being developed to design structures that deform into specified shapes given simple actuator inputs. In addition, these synthesis methods seek to optimize the stiffness of the structure to minimize actuator effort and maximize the stiffness with respect to the environment (external loading). These tools have been used to study two different types of morphing systems: (i) variable geometry wings and (ii) high-frequency vortex generators for active flow control. Several case studies are presented which highlight the design approach and computational and experimental results of these morphing aircraft systems.

Closed-Loop Missile Yaw Control via Manipulation of Forebody Flow Asymmetrics

A high-alpha, closed-loop flow-control system for missile yaw stabilization and enhanced maneuverability was designed, developed, and successfully demonstrated in a series of open- and closed-loop experiments on a fin-less 3:1 caliber tangent ogive missile model. The active flow-control-based yaw control system consisted of eight fastresponse pressure sensors and eight deployable flow effectors arranged in concentric rings on the missile nose cone. The devices were integrated with a closed-loop controller that modulated the effectors to manipulate flow asymmetry around the missile forebody. Side forces caused by crossflow separation and forebody flow asymmetries were observed on the missile model between 40 and 60 deg alpha. Parametric studies showed that actuating flow effectors in certain configurations resulted in cancellation of large side forces associated with the natural flow asymmetry under both steady and unsteady flow conditions. Exploratory studies conducted on optimized flow effector configurations resulted in control maps with wide spectrum of positive and negative side forces for yaw modulation. Dynamic experiments successfully demonstrated the ability of the closed-loop control system to generate and maintain a range of desired yawing moments during high-alpha pitch sweeps.

Active flow control using high-frequency compliant structures

Flow control to avoid or delay boundary-layer separation on a wing can dramatically improve the performance of most air vehicles in strategic parts of their individual flight envelopes. Previous aerodynamic experiments and computations have indicated that unsteady excitation at the appropriate frequency can delay boundary-layer separation and wing stall more effectively than steady flow perturbations and that these unsteady perturbations, when generated in an optimum frequency range, maximize the extent of flow separation control for specific flight conditions. Preliminary aerodynamic experiments have been performed on a deflected trailing-edge flap to evaluate turbulent boundary layer separation control with a deployable high-frequency micro-vortex-generator (HiMVG) array. The HiMVG design tested incorporated emerging displacement amplification compliant structures technology that deployed micro-vortex-generator blades 5 mm, through a range of frequencies between 30 and 70 Hz, when driven by an appropriately sized voice‐coil actuator. The mechanical HiMVG system tested produced an oscillatory stream of boundary-layer embedded vortices that proved effective in mitigating flow separation on the upper surface of a deflected flap when a similar array of static vortex generators could not. A second-generation HiMVG design driven by a piezoelectric actuator was also conceptualized. Candidate flow control applications for this second-generation design are discussed.

Closed-Loop Stall Control System

A closed-loop, stall sense and control system was demonstrated on a morphing airfoil. The FlexSys, Inc. Mission Adaptive Compliant Wing was modified to accept a Boeing Co. dielectric barrier discharge actuator panel in a location immediately upstream of the trailing-edge morphing flap, and hot-film sensors were installed on the model surface. A signal analysis algorithm, developed by Tao Systems, Inc., was applied to the hot-film signals to detect separation and trigger activation of the dielectric barrier discharge actuators. The system was successfully demonstrated in the U.S. Air Force Research Laboratory Phillip P. Antonatos Subsonic Aerodynamics Research Laboratory wind-tunnel facility, and an improvement in lift of about 10% was observed at Mach 0.05 (chord Reynolds number 9 x 10 5 ) under closed-loop control and a turbulent boundary-layer state. Actuator effectiveness was demonstrated up to Mach 0.1, but must be extended to Mach 0.2-0.3 to enable a practical stall control system for takeoff and approach of large aircraft. It may be possible to obtain that level of performance by optimizing the actuator locations and input waveforms.

Closed-Loop Stall Control on a Morphing Airfoil Using Hot-Film Sensors and DBD Actuators

A closed-loop, stall sense and control system was demonstrated on a morphing airfoil. The FlexSys, Inc. Mission Adaptive Compliant Wing (MACW) was modified to accept a Boeing Co. dielectric barrier discharge (DBD) actuator panel in a location immediately upstream of the trailing-edge morphing flap, and hot-film sensors were installed on the model surface. A signal analysis algorithm, developed by Tao Systems, Inc, was applied to the hot-film signals to detect separation and trigger activation of the DBD actuators. The system was successfully demonstrated in the AFRL SARL wind tunnel facility, and an improvement in lift of about 10% was observed at Mach 0.05 (chord Reynolds number 9× 10) under closed-loop control and a turbulent boundary layer state. Actuator effectiveness was demonstrated up to Mach 0.1, but must be extended to Mach 0.2–0.3 to enable a practical stall control system for takeoff and approach of large aircraft.

Structural Modal Control and Gust Load Alleviation for a SensorCraft Concept

SensorCraft is an Air Force Research Lab (AFRL) concept for a high flying vehicle that will be capable of providing greatly increased Intelligence, Surveillance, and Reconnaissance (ISR) capabilities. Part of the technology SensorCraft will require is high aerodynamic and structural efficiency to accomplish its mission goal of long range and sustained presence. To achieve a long loiter time, it will have a light weight structure with a high aspect ratio wing that carries much of the fuel mass. Hence, the structural modes will be in the same frequency range as the rigid body modes and will be strongly coupled with them. Wing stresses will need to be reduced and the gust load contribution will need to be minimized to keep the structural bending loads low along the span of the wing. This paper documents the results of a wind tunnel test at NASA Langley in the Transonic Dynamics Tunnel of a gust load alleviation system for a SensorCraft concept. The wing was fixed to the wall of the tunnel and the 4 trailing edge and 1 leading edge control effectors were used to simultaneously control first and second bending as well as control the pitching moment at the balance attachment.

SOME EXAMPLES OF AIRFOIL DESIGN FOR FUTURE UNMANNED AIR VEHICLE CONCEPTS

This paper presents design and analysis of three new airfoils for future unmanned air vehicles, operating at relatively low Reynolds numbers. The XFOIL and MSES computer codes were used to design, modify and analyze the airfoils. The first airfoil was designed for the joinedwing SensorCraft concept at Re = 10 6 , with the objective of maximizing section thickness and operation at supercritical Mach number. The second was designed for a new flying wing configuration with upper-surface blowing, operating at Re on the order of 200,000. The new airfoil has a large leading edge radius for ample plenum volume and a region of sharp pressure recovery immediately aft of the blowing slot. The third airfoil was designed for a notional fixed-wing micro air vehicle, operating at Re of 50,000, where the main objective was achieving a closed laminar separation bubble to obtain reasonable lift to drag ratios over a broach range of angle of attack. For the three baseline airfoils, the effects of Re, boundary layer transition parameter, and free vs. fixed transition were investigated. Further work is underway to validate the designs with experiments.

Supercritical Airfoil Design for Future High-Altitude Long-Endurance Concepts

The design and analysis is presented of a new laminar flow airfoil for a future high-altitude long-endurance aircraft that has an operational condition at supercritical speeds. The XFOIL and MSES computational codes were used to design, modify, and analyze the airfoil. The airfoil has enough thickness and performance to meet the requirements set for one of the U.S. Air Force Research Laboratorys SensorCraft concepts: a joined-wing configuration with a diamond shape in planform and front views. This SensorCraft concepts geometry and operational altitudes and speeds were used to determine the airfoil design conditions. The airfoil has a drag bucket over a large range of lift coefficients

SUPERCRITICAL AIRFOIL DESIGN FOR FUTURE HALE CONCEPTS

This paper presents design and analysis of a new airfoil for future High-Altitude Long-Endurance (HALE) aircraft that has an operational condition at supercritical speeds. The XFOIL and MSES computational codes were used to design, modify and analyze the airfoil. The airfoil has enough thickness and performance to meet the requirements set for one of the AFRL SensorCraft concepts; a joined-wing configuration with diamondshape in planform and front views. This SensorCraft concept’s geometry and operational altitudes and speeds were used to determine the airfoil design conditions. Sensitivity studies were carried out to investigate the effects of Reynolds number and Mach number, along with boundary layer transition parameters. The airfoil has a drag bucket over a large range of lift coefficient. Boundary layer transition location is at about 60% chord upper and 70% chord lower surface, and characterized by a laminar separation bubble, which decreases in size with increases in angle of attack. Further work needs to be performed to validate the design with experiments. NOMENCLATURE AR = Aspect ratio a = freestream speed of sound c = chord CL = airfoil lift coefficient CD = airfoil drag coefficient CM = airfoil moment coefficient CP = surface pressure coefficient L = lift force D = drag force M = freestream Mach number, V/a MD = drag divergence Mach number M = reduced Mach number, L

An Adaptive Structures Electro -Mechanical Device for Dynamic Flow Control Applications

Flow control to prevent or delay boundary layer separation can dramatically improve the performance of air vehicles in critical regions of the flight envelope. Prior aerodynamic experim ents have shown that unsteady excitation, at the appropriate frequency, can delay boundary layer separation more effectively than steady flow perturbations. An electro - mechanical flow control device, with a substantial deployment frequency bandwidth, has been developed and tested in both static and dynamic flow control environments. The device couples a high frequency piezostack actuator with a sixty five -to -one displacement amplification mechanism to oscillate sixteen vortex generating blades at rates up t o two hundred hertz. Conceptual design of the piezostack -compliant mechanism is discussed. Flow separation control results are presented and device performance issues including dynamic characteristics are addressed.