Showing papers on "Pitching moment published in 2006"

Rotor Airloads Prediction Using Loose Aerodynamic/Structural Coupling

Abstract: A computational fluid dynamics (CFD) code and rotorcraft computational structural dynamics (CSD) code are coupled to calculate helicopter rotor airloads across a range of flight conditions. An iterative loose (weak) coupling methodology is used to couple the CFD and CSD codes on a per revolution, periodic basis. The CFD code uses a high fidelity, Navier‐Stokes, overset grid methodology with first principles-based wake capturing. Modifications are made to the CFD code for the aeroelastic analysis. For a UH-60A Blackhawk helicopter, three challenging level flight conditions are computed: 1) high speed, μ = 0.37, with advancing blade negative lift, 2) low speed, μ = 0.15, with blade‐vortex interaction, and 3) high thrust with dynamic stall, μ = 0.24. Results are compared with UH-60A Airloads Program flight test data. For all cases the loose coupling methodology is shown to be stable, convergent, and robust with full coupling of normal force, pitching moment, and chord force. In comparison with flight test data, normal force and pitching moment phase and magnitude are in good agreement. The shapes of the airloads curves are well captured. Overall, the results are a noteworthy improvement over lifting line aerodynamics used in rotorcraft comprehensive codes.

Dynamic stall modelling of the S809 aerofoil and comparison with experiments

Abstract: An analysis of dynamic stall for the S809 aerofoil has been performed in conjunction with the Leishman–Beddoes dynamic stall model that was modified for wind turbine applications. Numerical predictions of the lift, drag and pitching moment coefficients were compared with measurements obtained for an oscillating S809 aerofoil at various reduced frequencies, mean angles of attack and angle of attack amplitudes. It was found that the results using the modified model were in good agreement with the experimental data. Hysteresis in the aerodynamic coefficients was captured well, although the drag coefficient was slightly underpredicted in the deep stall flow regime. Validation against the experimental data showed overall good agreement.The mathematical structure of the model is such that it can be readily incorporated into a comprehensive analysis code for a wind turbine. Copyright © 2006 John Wiley & Sons, Ltd.

Dynamic Stall Flow Control via a Trailing-Edge Flap

Abstract: Effects of a moveable trailing-edge flap on the dynamic-load loops of an oscillating NACA 0015 airfoil were investigated at Re =1 .65 × × 10 5 . Both upward and downward flap deflections for different flap actuation start times, durations, and amplitudes were considered. Surface pressure measurements show that only upward flap deflections led to the reduction of nose-down pitching moment, and that the larger the flap deflection the more efficient the reduction mechanism was. The shorter the actuation duration the smaller the poststall lift loss. The reduction of the peak nose-down pitching moment was a consequence of the suction pressure appearing on the lower surface of the trailing-edge flap. The formation and detachment of the dynamic-stall vortex was not affected by the flap motion, whereas the low-pressure signature of the vortex was reduced by the upward flap deflection. The later the flap actuation during the upstroke, the greater the net torsional damping was. Optimum flap motion is also discussed.

Non-intrusive load characterization of an airfoil using PIV

Abstract: An assessment is made of the feasibility of using PIV velocity data for the non-intrusive aerodynamic force characterization (lift, drag and pitching moment) of an airfoil. The method relies upon the application of control-volume approaches in combination with the deduction of the pressure from the PIV experimental data, by making use of the momentum equation. First, the consistency of the method is verified by means of synthetic data obtained from CFD. Subsequently, the procedure was applied in an experimental investigation, in which the PIV approach is validated against standard pressure-based methods (surface pressure distribution and wake rake).

Multidisciplinary design optimization of wing shape for a small jet aircraft using kriging model

Abstract: *† ‡ § ¶ # In this paper, a practical Multidisciplinary Design Optimization (MDO) system for an aircraft design is developed The MDO system is based on the integration of computational fluid dynamics (CFD) codes and the NASTRAN based aeroelastic-structural interface code Kriging model is employed to save the computational time of objective function evaluation in Multi-Objective Genetic Algorithm (MOGA) As a result of optimization, several nondominated solutions, indicating the trade-off among the drag, the structural weight, the drag divergence and the pitching moment, have been found

The Application of Variable Cant Angle Winglets for Morphing Aircraft Control

Abstract: This paper investigates a novel method of control for ‘morphing’ aircraft. The concept consists of a pair of winglets with adaptive cant angle, independently actuated, mounted at the tips of a flying wing. Computations with a vortex lattice model and wind tunnel tests demonstrate the validity of the concept. The variable cant angle winglet appears to be a multi-axis effector with a favorable coupling in pitch and roll with regard to turning manoeuvres. Nomenclature � angle of attack �e elevator deflection angle � reduced spanwise coordinate vortex strength l, r left and right winglet cant angles x vortex strength sensitivity w.r.t. variable x turn rate � bank angle � air density b wing span CD, CY , CL drag, side-force and lift coefficients Cl, Cm, Cn rolling, pitching and yawing moment coefficients CDx , CYx , CLx drag, side-force and lift coefficient derivatives w.r.t. parameter x Clx , Cmx , Cnx rolling, pitching and yawing moment coefficient derivativesw.r.t. parameter x croot root chord g gravitational acceleration p, q, r aircraft rotation rates in body or stability axes R turn radius u, v, w aircraft velocity components in body or stabilty axes Vcg flight speed W aircraft weight CG center of gravity LE leading edge TE trailing edge

Aerodynamic loads on airfoil with trailing-edge flap pitching with different frequencies

Abstract: Unsteady aerodynamic loads on NACA 0012 airfoil with a trailing-edge flap were measured in wind tunnel and calculated from a simple theoretical model. The airfoil model of 0.18 m chord length used in wind-tunnel test was oscillating in pitch about an axis located at 35% chord length from the airfoil leading edge. The length of trailingedge flap was 22.6% of airfoil chord. The flap was also deflecting harmonically, but with frequency different from airfoil pitching motion. The influence of phase delay between airfoil angle of incidence and flap deflection at the beginning of the motion was considered. The theoretical method used for calculation of unsteady airfoil loads is based on two-dimensional, inviscid, incompressible flow model at subsonic Mach numbers. The expressions for unsteady aerodynamic loads calculations on the airfoil and on the flap were obtained in a closed form using distribution of flow velocity potential along the airfoil chord and along the flap length. Lift and aerodynamic moment measured in the wind tunnel were compared with results of calculations. The correlation between experimental and theoretical results is adequate.

Design of Flapping Airfoil for Optimal Aerodynamic Performance in Low-Reynolds Number Flows

Abstract: The authors appreciate the financial support by the Basic Research Program of the Korea Science & Engineering Foundation (Grant No. R01-2002-000-00329-0), and by the Brain Korea-21 Program for the Mechanical and Aerospace Engineering at Seoul National University.

Control of rotorcraft retreating blade stall using air-jet vortex generators

Abstract: A series of low-speed wind tunnel tests was carried out on an oscillating airfoil fitted with two rows of air-jet vortex generators (AJVGs) The airfoil used had an Royal Aircraft Establishment 9645 section, and the two spanwise arrays of AJVGs were located at x/c = 012 and 062 The devices and their distributions were chosen to assess their ability to modify/control dynamic stall, the goal being to enhance the aerodynamic performance of helicopter rotors on the retreating blade side of the disk The model was pitched about the quarter chord with a reduced frequency k of 01 in a sinusoidal motion defined by a = 15 + 10sin ωt deg The measured data indicate that, for continuous blowing from the front row of AJVGs with a momentum blowing coefficient C μ greater than 0008, modifications to the stalling process are encouraging In particular, the pitching moment behavior exhibits delayed stall and there is a marked reduction in the normal force hysteresis

Aerodynamic Design Studies of Conventional & Unconventional Wings with Winglets

Abstract: In the pursuit of efficiency there is continued and developing interest in the subject of winglets and novel designs such as Blended Wing Body and Oblique Flying Wing. Winglets may exist in many modes, ranging from endplates, through folded wing tips to complex upper and lower surface partial chord devices. We have developed a design process that is directly applicable to the design of wings and winglet combinations. A set of target conditions is established. These may be, for example, minimum liftinduced drag, minimum wave drag, a known structural loading or pitching moment control. The design process is iterative and, in general, is driven towards the idealised spanwise load distribution for minimum drag. A degree of pitch control has also been demonstrated on selected configurations. The method may be used to design from “scratch” or in the corrective sense to optimise a current design for new requirements at the design point. The method has been applied to the design of several configurations of current interest, conventional A340 type wing, Blended Wing Body and Oblique Flying Wing. Winglets have been simulated by folding the outer wing through various angles. The technique has proved to be easy to use. It is enlightening as it gives, at every stage, a feel for what is happening in terms of camber development, pressure distributions and Centre of Pressure location. Favourable characteristics of the configuration can be enhanced whereas those that are not beneficial can be minimised or avoided as the design progresses.

Simultaneous Pseudo-Timestepping for Aerodynamic Shape Optimization Problems with State Constraints

Abstract: This paper presents a numerical method for constrained aerodynamic shape optimization problems. It is based on simultaneous pseudo-timestepping in which stationary states are obtained by solving the pseudo-stationary system of equations representing the state, costate, and design equations. The main advantages of this method are that it blends in nicely with a previously existing pseudo-timestepping method for the states only, and that a preconditioner can be used for convergence acceleration which stems from the reduced SQP methods. Design examples for drag reduction with constant lift and drag reduction with constant pitching moment for an RAE2822 airfoil are included. The overall cost of computation is 4--8 times that of the forward simulation runs depending on the grid size and the number of design parameters.

Oscillating Wing Loadings with Trailing-Edge Strips

Abstract: The dynamic-load loops of an oscillating NACA 0012 airfoil, obtained from surface-pressure measurements, fitted with small full-span trailing-edge strips, positioned on the lower (i.e., the Gurney flaps) and/or upper (the inverted strips) surface at the wing trailing edge, of height of 1.6 and 3.2% of the airfoil chord were studied at Re = 1.07 x 105. The results show that, similar to a static wing, the Gurney flap concept was also fairly generally applicable, in terms of lift performance enhancement, to an oscillating airfoil, despite the large and small increases in the peak negative pitching moment and maximum drag force, respectively, as well as a promoted dynamic stall. The increase (decrease) in the nose-down pitching moment (dynamic-stall angle) can be alleviated by the use of inverted strips at the price of a reduced maximum lift coefficient. The asymmetric strips were found to provide a compromise in the dynamic aerodynamic performance between the regular and inverted Gurney flaps, including a 49-deg flap, of an oscillating airfoil. The present passive oscillating-wing C 1 -C d -C m control can be valuable, because it can be used as an experimental guideline for the active control of the dynamic stall and/or nose-down pitching moment via a spanwise trailing-edge flap.

Scramjet Lift, Thrust and Pitching-Moment Characteristics Measured in a Shock Tunnel

Abstract: Lift, pitching moment, and thrust/drag on a supersonic combustion ramjet were measured in the T4 free-piston shock tunnel using a three-component stress-wave force balance. The scramjet model was 0.567 m long and weighed approximately 6 kg. Combustion occurred at a nozzle-supply enthalpy of 3.3 MJ/kg and nozzle-supply pressure of 32 MPa at Mach 6.6 for equivalence ratios up to 1.4. The force coefficients varied approximately linearly with equivalence ratio. The location of the center of pressure changed by 10% of the chord of the model over the range of equivalence ratios tested. Lift and pitching-moment coefficients remained constant when the nozzle-supply enthalpy was increased to 4.9 MJ/kg at an equivalence ratio of 0.8, but the thrust coefficient decreased rapidly. When the nozzle-supply pressure was reduced at a nozzle-supply enthalpy of 3.3 MJ/kg and an equivalence ratio of 0.8, the combustion-generated increment of lift and thrust was maintained at 26 MPa, but disappeared at 16 MPa. Measured lift and thrust forces agreed well with calculations made using a simplified force prediction model, but the measured pitching moment substantially exceeded predictions. Choking occurred at nozzle-supply enthalpies of less than 3.0 MJ/kg with an equivalence ratio of 0.8. The tests failed to yield a positive thrust because of the skin-friction drag that accounted for up to 50% of the fuel-off drag.

Ducted Rotor Performance Calculations and Comparisons with Experimental Data

Abstract: *† ‡ Ducted fan performance was calculated using the AMI panel method VSAERO coupled with the ROTOR propeller module. Calculations were performed for a set of experimental data measured on a test model in the NASA Ames facility. The ducted fan model was a generic representation of a class of proposed small UAV systems. Good correlation of duct and rotor thrust and aerodynamic pitching moments was found for all conditions under which the duct lip flow was attached. Empirical extensions to stalled duct lip results can be made for high rotor thrust. I. Introduction new class of unmanned aerial vehicles under development are small ducted fans. These vehicles have no separate lifting surfaces but instead rely on rotor and fan-induced thrust for lift and maneuvering. These UAVs operate primarily at low speeds, high thrust, and high angles of attack. An area of particular concern for these UAVs is translational performance in hover. In this flight condition (nearly 90 degrees angle of attack with respect to the duct thrust axis) a large aerodynamic pitching moment is generated which tends to pitch the vehicle away from the direction of translation. This pitching moment requires additional control authority to overcome and in many cases can severely limit the translational speed of the hovering vehicle. A computer code well suited to the task of calculating these flows is the ROTOR module of the AMI panel method VSAERO. ROTOR computes the coupled effects of the propeller induced flow and the external flow around the duct. The code combines a blade element theory propeller calculation with the VSAERO potential flow solution for arbitrary bodies in subsonic flow. The code takes into account induced effects of the propeller such as asymmetric inlet lip suction and the effects of the exhaust flow in thrust vectoring. Computations for a generic ducted fan UAV compare favorably with test results. The overall lift, drag, and pitching moments developed by the vehicle can be calculated with the VSAERO/ROTOR method for all flight conditions in which attached flow is maintained around and through the duct inlet. This includes the high angle of attack and high thrust conditions (representing crosswinds in hover) which are of particular concern to this class of vehicle. While the very high duct angles of attack would initially seem to preclude the successful use of a potential flow method, the high duct thrust and resulting high mass flows through the duct maintain attached flow through the duct. The resulting flow field can be successfully calculated with this method. . Empirical correlations of duct stall can be used to provide limits of attached flow behavior and guidelines for estimates of aerodynamic forces and moments beyond duct stall.

Combining passive control method for dynamic stall control

Abstract: To demonstrate the benefits of a combined nose droop and Gurney flap for improving dynamic stall and poststall aerodynamic characteristics of a rotor airfoil, numerical investigations and design optimization have been performed. As a means of passive flow control, a fixed nose droop has been deliberately employed together with a Gurney flap. For shape optimization, droop location and droop angle were selected as design variables for the fixed nose droop, and the flap length was chosen as a design variable for the Gurney flap. Bousman's function plot was employed to define the objective function and constraint conditions. A feasible direction-based optimizer and a higher-order response surface method were harnessed to handle the highly nonlinear properties of dynamic stall. By the use of this methodology, optimum design was carried out to enhance lift and pitching moment characteristics simultaneously; at.the same time, the unfavorable effects of the overall movement of the pitching moment coefficient induced by droop and the Gurney flap was reduced. It is also proved that by utilizing a 22.7-deg droop at the 0.275 chord droop position and a 1.14% chord Gurney flap, the stall can be delayed by a maximum lift coefficient increased by 13%, a maximum negative pitching moment reduced by 60%, and the lift-to-drag ratio increased accordingly. The present combined passive control methods and design result show significant improvement of aerodynamic performance in Bousman's plot in terms of lift and pitching moment.

Adaptive Control Using Neural Network Augmentation for a Modified F-15 Aircraft

Abstract: Description of the performance of a simplified dynamic inversion controller with neural network augmentation follows. Simulation studies focus on the results with and without neural network adaptation through the use of an F-15 aircraft simulator that has been modified to include canards. Simulated control law performance with a surface failure, in addition to an aerodynamic failure, is presented. The aircraft, with adaptation, attempts to minimize the inertial cross-coupling effect of the failure (a control derivative anomaly associated with a jammed control surface). The dynamic inversion controller calculates necessary surface commands to achieve desired rates. The dynamic inversion controller uses approximate short period and roll axis dynamics. The yaw axis controller is a sideslip rate command system. Methods are described to reduce the cross-coupling effect and maintain adequate tracking errors for control surface failures. The aerodynamic failure destabilizes the pitching moment due to angle of attack. The results show that control of the aircraft with the neural networks is easier (more damped) than without the neural networks. Simulation results show neural network augmentation of the controller improves performance with aerodynamic and control surface failures in terms of tracking error and cross-coupling reduction.

Morphing Airfoil Design for Minimum Aerodynamic Drag and Actuation Energy Including Aerodynamic Work

Abstract: Recently, there has been great interest in developing technologies that may enable a morphing aircraft. Such an aircraft can change shape in flight, which would make it possible to adjust the wing to the best possible shape for any flight condition encountered by the aircraft. However, there is an actuation cost associated with making these shape changes that must be included in the optimization process. A previous effort investigated a simple strain energy model to account for the actuation cost for a morphing airfoil, and a multiobjective optimization found tradeoff solutions between low-energy, high-drag and highenergy, low-drag morphing airfoils. Building upon the previous effort, the main purpose of this paper is to formulate the aerodynamic contribution to work on the morphing airfoil. This aerodynamic work term can be added to the strain energy model to compute the total energy required for changing the shape of a morphing airfoil. With this formulation, a beneficial contribution of aerodynamic work could reduce the morphing actuation energy from that associated solely with deformation of the aircraft structure, or an unfavorable contribution of aerodynamic work could increase the morphing actuation energy. Studies presented in this paper illustrate how the aerodynamic work is computed and how the aerodynamic work can be either beneficial or unfavorable. The effect of the morphing airfoil’s relative stiffness on the multi-objective solutions is also presented. Nomenclature i d C = Coefficient of drag at design condition i i l C = Coefficient of lift at design condition i m C = Coefficient of pitching moment f C = Coefficient of skin friction p C = Coefficient of pressure c = Airfoil chord length e K = Unit vector from State 1 to State 2 a f = Actuator force s f = Internal force in spring

Performance predictions of the nrel phase vi combined experiment rotor using a free-vortex wake model

Abstract: This paper describes the application of a free-vortex wake method (FVM) to predict the performance and aerodynamic loads on an upwind wind turbine in unyawed and yawed flow. Measurements from the Phase VI of the NREL/NASA Ames wind tunnel test were used for validating the predictions. The FVM predicted well the variation of the turbine thrust and the aerodynamic power output with wind speed. Spanwise distributions of the aerodynamic coefficients (normal force, leading-edge thrust and pitching moment) were also predicted, and encouraging agreement was obtained against the measured coefficients. The azimuthal variation of loads showed that the unsteady aerodynamic behavior of the the wind turbine was predicted well, with some exceptions. However, the vibration modes of the blade, tower, and drive-train seemed to have a significant influence on the measured loads in the experiment.

Aerodynamic Analysis of Grid-Fin Configurations at Supersonic Speeds

Abstract: An aerodynamic prediction method for the estimation of aerodynamic characteristics of grid-fin configurations at supersonic Mach numbers has been developed. The method is based on shock-expansion theory. The shock and expansion relations and the interactions between the shock and expansion waves are used to predict the pressure distribution inside each grid-fin cell. The effect of body is modeled using doublet in the crossflow plane. The effect of symmetric separated vortices from the leeward side of body is also modeled using empirical data for the strength and location of vortices. The present method has been validated with available experimental data for body grid-fin configurations at angles of attack without and with fin deflection. The comparison of normal force coefficient on the individual grid fins as well as the overall normal force, pitching moment, and axial force coefficient with experimental data is good. Effect of roll orientation on normal force, pitching moment, and induced out-of-plane force is also brought out.

Cf bearing with steady pitching moment

Abstract: A centrifugal force bearing having a means for providing a steady pitching moment is disclosed. The centrifugal force bearing may optionally comprise a coning means. A rotor system having the centrifugal force bearing is disclosed. A rotary-wing aircraft having the centrifugal force bearing is disclosed.

Parameter Estimation from Flight Data of a missile using Maximum Likelihood and Neural Network Method

Abstract: [] Parameter estimation from flight data as applied to aircraft, missile in the linear flight regime is currently being used on routine basis. However, the linear aerodynamic models used successfully upto this time seem to be inadequate for newly introduced short range, highly maneuverable missile, rockets etc. The aerodynamics is required throughout the design process of any flight vehicle. Rapid progress has been made in the use of the so-called output error method, such as the modified Newton-Raphson method and maximum likelihood (ML) methods. Several requirements have been introduced to improve accuracy and reliability of estimates. The presence of measurement and process noise and this, in general, has resulted in relatively increased complexity and computational time. Application of these methods requires a priori postulations of the aerodynamic model. The application of the Delta method and ML method on flight data of a typical tactical missile faces one main difficulty. Due to operational reasons, it might not be possible to excite the missile with a multistep efficient control input and that might result in having flight data with inadequate information content for the purpose of parameter estimation. The present study makes an attempt to estimate aerodynamic parameters from a typical flight data of a short range missile. Since ML method requires a priori postulation of the model, exhaustive wind tunnel testing were conducted to generate longitudinal force and moment coefficients. Identification methods were applied on the selected wind tunnel data to capture the general form of the aerodynamic model. During the application of ML method, the estimation algorithm assumed this wind tunnel identified aerodynamic model to be exact. To avoid any requirement of postulation of aerodynamic model, the Delta method was applied to estimate the aerodynamic parameters. The Delta method used measured aircraft motion and control variables as the inputs to the Feed Forward Neural Network and the aerodynamic force or moment coefficient was the output for training the Feed Forward Neural Network. The application of the Delta method results in large scatters in the estimated parameters. To overcome this problem of large scatter, the Delta method was modified by changing the training strategy. The Delta method with new training strategy will be referred as the Modified Delta method. It is expected that the proposed Modified Delta method would result in estimates with less uncertainties. Further to check the robustness of the both ML and Delta as the Modified delta methods, the estimation was also carried out with flight data having known measurement noise. The effect of control input form in the accuracy of estimates obtained by ML and the Delta as the Modified Delta methods are also studied. It is observed that the Modified Delta method can advantageously be applied on the flight data of a tactical missile to estimate aerodynamic parameters. The paper progresses with the description of the generation of wind tunnel data and aerodynamic model identification using selected wind tunnel data. Finally it concludes by demonstrating applicability of ML and the Delta as the Modified Delta methods on simulated flight data of a typical short range tactical missile configuration.

Gurney Flaps on Slender and Nonslender Delta Wings

Abstract: *† ‡ A parametric low-speed wind tunnel study has been undertaken of the effect of Gurney flaps on the aerodynamic characteristics of slender and non-slender delta wings with sweep angles of 40°, 60° and 70°. In contrast to published data, no fundamental difference in behaviour with sweep angle was found. Gurney flaps were found to give similar lift and pitching moment changes to conventional hinged control surfaces, but with a much greater drag penalty. Gurney flap effectiveness was found to increase as sweep angle is reduced; a characteristic explained by a simple analysis of lift and drag on a sharp-edged delta wing. For a 40° sweep angle (typical of a UCAV configuration), Gurney flap deployment can significantly improve lift/drag ratio at higher incidences. Further improvements can be obtained by tapering or serrating the flap.

Flight dynamics of the screw kick in rugby

Abstract: This paper describes the aerodynamic forces and the flight trajectory for the screw (spiral) kick in rugby The screw kick is defined as that which causes the ball to spin on its longitudinal axis The aerodynamic forces acting on a rugby ball spinning on its longitudinal axis were measured in a wind tunnel using a six-component strut type balance It was found that the drag, the lift and the pitching moment depend on the angle of attack, while the side force (Magnus force) depends on both the spin rate and the angle of attack in the range where the wind speed lies between 15 and 30 m s-1 and the spin rate is between 1 and 10 revolutions per second Moreover, the flight trajectory was obtained by integrating the full nonlinear six degrees of freedom equations of motion on the basis of aerodynamic data In a simulation, a ball spinning on its longitudinal axis tended to hook toward or away from the touchline even if the velocity and angular velocity vectors were parallel to the touchline The direction of the hook depends on the direction of the angular velocity vector The initial direction of the hook depends on the relationship between the flight path angle and the pitch angle as well as the direction of the angular velocity vector

Numerical Simulation of the AEDC Waverider at Mach 8

Abstract: A waverider designed for Mach 14 flight is numerically simulated using the University of Minnesota US3D computational fluid dynamics code. Numerical results are compared with experimental data provided by the Arnold Engineering Development Center Hypervelocity Wind Tunnel Number 9. Two sets of nominal Mach 8 flight conditions are examined, with a low Reynolds number of 14 million/m and a high Reynolds number of 54 million/m. Pitch angle sweeps in the range of -10° to 5° are presented for both cases. Very good agreement between numerical simulation and experimental measurement is shown in plots of lift-todrag ratio vs. alpha, lift and drag coefficients vs. alpha, and pitch moment coefficient vs. alpha. In addition, direct comparisons are made for the complete set of discrete pressure taps and heat transfer probes.

Short duration force measurements in impulse facilities

Abstract: A new multiple component stress wave force balance has been designed, calibrated and tested in the High Enthalpy Shock Tunnel Gottingen (HEG) of the German Aerospace Center (DLR). The balance is able to measure forces of short duration (milliseconds) on instrumented models from angles of attack from -40 to 20 degrees. Two models, a blunt cone 303mm long and a standard force reference model (HB-2) of 70mm diameter, are used to establish the accuracy of the force balance. The blunt cone tests were conducted at two different test conditions with a constant Mach number of 7.8 and total enthalpies of 3.0 and 3.5MJ/kg. At 0 degrees angle of attack and an enthalpy of 3.0MJ/kg, the measured axial coefficient was recovered to within 6% when compared to computational fluid dynamic (CFD) simulations. At –10 degrees, the axial and normal coefficients were within 6% and 9% respectively of CFD predictions while the center of pressure (based on chord length) was within 2%. Tests with the HB-2 standard force reference model were conducted at enthalpies ranging from 12 to 22 MJ/kg at an angle of attack of 0 degrees. A linear variation of the axial coefficient with the viscous similarity parameter was predicted with non-equilibrium CFD simulations assuming a laminar boundary layer. The recovered axial force coefficient remained within 6% of the CFD predictions and compared well with experimental results from other wind tunnel facilities. Reasonable comparison of pressure and heat flux measurements along longitudinal symmetry lines of the model was obtained. The accuracy of the force balance is estimated at approximately 5% for the axial component and 4% for the normal and pitching moment components.

Assessment of an Unstructured-Grid Method for Predicting Aerodynamic Performance of Jet Flaps

Abstract: The application of a Computational Fluid Dynamics tool to a jet flap control effector on an elliptical airfoil-section wing was investigated. The study utilized the Tetrahedral Unstructured Software System developed at NASA Langley Research Center. The Reynolds-averaged Navier-Stokes flow solver code used was USM3D. The CFD-based jet flap simulations were compared to experimental results from a wind tunnel test conducted at the NASA Langley Transonic Dynamics Tunnel. The wind tunnel model consisted of a six percent thick elliptical airfoil with a modified trailing edge. The jet flap was located at 95% chord and exited at 90 degrees to the lower surface. The experimental model was designed to promote two-dimensional flow across the wing. It was found that the CFD simulation had to model the three-dimensional geometry of the experiment in order to obtain good agreement. Tests were performed at two Mach numbers at several different jet momentum coefficients. In order to be consistent with the experimental method, the CFD lift and pitching moment values were determined by integrating the pressures over the wing.