David A. Seidel

Time-marching transonic flutter solutions including angle-of-attack effects

Transonic aeroelastic solutions based upon the transonic small perturbation potential equation were studied. Time-marching transient solutions of plunging and pitching airfoils were analyzed using a complex exponential modal identification technique, and seven alternative integration techniques for the structural equations were evaluated. The HYTRAN2 code was used to determine transonic flutter boundaries versus Mach number and angle-of-attack for NACA 64A010 and MBB A-3 airfoils. In the code, a monotone differencing method, which eliminates leading edge expansion shocks, is used to solve the potential equation. When the effect of static pitching moment upon the angle-of-attack is included, the MBB A-3 airfoil can have multiple flutter speeds at a given Mach number.

Unsteady transonic flow calculations for realistic aircraft configurations

A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.

Experimental flutter boundaries with unsteady pressure distributions for the NACA 0012 Benchmark Model

The Structural Dynamics Division at NASA Langley Research Center has started a wind tunnel activity referred to as the Benchmark Models Program. The primary objective of the program is to acquire test data that will be useful for developing and evaluating aeroelastic type CFD codes currently in use of under development. This paper describes the progress achieved in testing the first model in the Benchmark Models Program. Experimental flutter boundaries are presented for a rigid semispan model (NACA 0012 airfoil section) mounted on a flexible mount system. Also, steady and unsteady pressure measurements taken at the flutter condition are presented. The pressure data were acquired over the entire model chord located at the 60 percent span station.

Recent advances in transonic computational aeroelasticity

Abstract A transonic unsteady aerodynamic and aeroelasticity code called CAP-TSD has been developed for application to realistic aircraft configurations. The code permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis in the flutter critical transonic speed range. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance potential equation. The paper gives an overview of the CAP-TSD code development effort and presents results which demonstrate various capabilities of the code. Calculations are presented for several configurations including the General Dynamics one-ninth scale F-16C aircraft model and the ONERA M6 wing. Calculations are also presented from a flutter analysis of a 45° sweptback wing which agree well with the experimental data. The paper presents descriptions of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate these recent developments in transonic computational aeroelasticity.

Transonic stability and control of aircraft using CFD methods

Implementation of a capability to calculate longitudinal short-period response in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) finite-difference code is described. The code, developed recently at the NASA Langley Research Center, is capable of solving steady and unsteady flows about complete aircraft configurations and is used primarily for aeroelastic calculations in the critical transonic speed range. The longitudinal short-period equations of motion in state-space form have been coupled to the time-accurate lift and moment calculated by the program. Transient responses to an elevator pulse for free-flying aircraft demonstrate the new capability. A trim routine is also added to the code to obtain trim automatically during steady-state flow field convergence. Stability and control derivatives are estimated from the calculated transient response by a maximum likelihood estimation program. Results for a fighter configuration and a general aviation configuration are presented to assess the capability.

Transonic shock-induced dynamics of a flexible wing with a thick circular-arc airfoil

Transonic shock boundary layer oscillations occur on rigid models over a small range of Mach numbers on thick circular-arc airfoils. Extensive tests and analyses of this phenomena have been made in the past but essentially all of them were for rigid models. A simple flexible wing model with an 18 pct. circular arc airfoil was constructed and tested in the Langley Transonic Dynamics Tunnel to study the dynamic characteristics that a wing might have under these circumstances. In the region of shock boundary layer oscillations, buffeting of the first bending mode was obtained. This mode was well separated in frequency from the shock boundary layer oscillations. A limit cycle oscillation was also measured in a third bending like mode, involving wind vertical bending and splitter plate motion, which was in the frequency range of the shock boundary layer oscillations. Several model configurations were tested, and a few potential fixes were investigated.

Transonic region of high dynamic response encountered on an elastic supercritical wing

Unsteady aerodynamic data were measured on a 10.3 aspect-ratio elastic supercritical wing while undergoing high dynamic response above a Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based upon subcritical response data. During the present test, no instability was found, but an angle-of-attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outboard control surface deflection and a lower surface spanwise fence located near the 60% local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Transonic unsteady pressure measurements on a supercritical airfoil at high Reynolds numbers

Steady and unsteady pressures were measured on a two-dimensional supercritical airfoil in the NASA Langley Research Center 0.3 m Transonic Cryogenic Tunnel at Reynolds numbers of 6 to 35xl06. The airfoil was oscillated in pitch at amplitudes of ±0.25 to ±1.0 deg at frequencies of 5-60 Hz. The special requirements of testing an unsteady pressure model in a pressurized cryogenic tunnel are discussed. Selected steady measured data are presented and are compared with GRUMFOIL calculations at Reynolds numbers of 6 and 30 x 106. Experimental unsteady results at Reynolds numbers of 6 and 30 X106 are examined for Reynolds number effects. Measured unsteady results at two mean angles of attack at a Reynolds number of 30 x 106 are also examined. Nomenclature Cp = pressure coefficient \CP I = modulus of oscillating pressure coefficient Cj = lift coefficient c = chord, in. / = frequency, Hz k = reduced frequency, based on semichord, ircf/V M = Mach number R = Reynolds number based on chord V — velocity, in./s x = stream wise coordinate measured from leading edge, in. a =peak oscillation amplitude in pitch, positive leading edge up, deg a = steady or mean angle of attack, positive leading edge up, deg <£ = phase angle between oscillating pressure and oscillating wing angle of attack, deg

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above a Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based on subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outbound control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Highlights of unsteady pressure tests on a 14 percent supercritical airfoil at high Reynolds number, transonic condition

Steady and unsteady pressures were measured on a 2-D supercritical airfoil in the Langley Research Center 0.3-m Transonic Cryogenic Tunnel at Reynolds numbers from 6 x 1,000,000 to 35 x 1,000,000. The airfoil was oscillated in pitch at amplitudes from plus or minus .25 degrees to plus or minus 1.0 degrees at frequencies from 5 Hz to 60 Hz. The special requirements of testing an unsteady pressure model in a pressurized cryogenic tunnel are discussed. Selected steady measured data are presented and are compared with GRUMFOIL calculations at Reynolds number of 6 x 1,000,000 and 30 x 1,000,000. Experimental unsteady results at Reynolds numbers of 6 x 1,000,000 and 30 x 1,000,000 are examined for Reynolds number effects. Measured unsteady results at two mean angles of attack at a Reynolds number of 30 x 1,000,000 are also examined.