Transonic Effects Research Articles

Trans-phonon effects in ultra-fast nanodevices

We report a novel phenomenon in carbon nanotube based ultra-fast mechanical devices, thetrans-phonon effect, which resembles the transonic effects in aerodynamics. It is causedby dissipative resonance of nanotube phonons similar to the radial breathingmode, and subsequent drastic surge of the dragging force on the sliding tube,and multiple phonon barriers are encountered as the intertube sliding velocityreaches critical values. It is found that the trans-phonon effects can be tuned byapplying geometric constraints or varying chirality combinations of the nanotubes.

Navier-Stokes Analysis of Helicopter Rotor Aerodynamics in Hover and Forward Flight

Reynolds-averaged compressible Navier ‐Stokes computations are presented for the hovering 7A model rotor and a low aspect ratio NACA0012 rotor in nonlifting forward e ight. For enhanced hover performance prediction, aeroelastic effects are taken into account by tightly coupling the e ow solver with a e nite element model of the blade based on Timoshenko beam theory. Good agreement between computed and measured rotor performance is achieved for three collective pitch settings on structured periodic grids featuring noncongruent cell faces along the periodicity planes. Results of a hybrid Navier ‐Stokes/Euler chimera hover computation are presented, in addition to the periodic grid analysis demonstrating the superiority of the overset grid approach with respect to tip vortex conservation. The unsteady Navier ‐Stokes computations for the nonlifting forward e ight test case show more sensitivity relative to time step size than comparable Euler analyses. Overall correlation of computed and experimental near-tip pressure distributions over blade azimuth is considered to be fair, and the strong transonic effects on the advancing blade are adequately captured by the numerical analysis.

NUMERICAL INVESTIGATIONS OF THE AERODYNAMICS AND THE AEROELASTIC STABILITY OF OSCILLATING ANNULAR WINGS

The nacelles of modern aeroengines are constantly increasing in size. Thus, engine air-loads are becoming more powerful and their importance for the aeroelastic stability is becoming more significant. The principal goal of this study is to answer the question of how unsteady airloads vary while shifting to transonic Mach numbers. The investigations are carried out by applying a finite volume Euler method to a harmonically oscillating annular wing. The results show that transonic effects in the case of an annular wing are essentially weaker than in the case of an airfoil. The order of magnitude of the variations is around 10%. Possible consequences for the aeroelastic stability are examined with the example of an elastically mounted annular wing in transonic flow. The shifts of the stability curves also remain within a range of 10%. In addition, an actuator disk method, which is frequently used for the simulation of the fan jet, is expanded in such a way that unsteady flows can be treated. Some unsteady air-loads are strongly dependent on the pressure jump across the fan.

The main sources of helicopter vibration and noise emissions and adaptive concepts to reduce them

AbstractThis paper deals with the sources of helicopter vibration and noise emissions. It shows and evaluates adaptive concepts to reduce them. The main reasons for these emissions are aerodynamic effects such as blade‐vortex‐inter action (BVI), aerodynamic hysteresis, and local transonic effects at the forward moving blade.After a review of conventional control concepts to reduce the vibrations and/or sound emissions, the possibilities to realise an adaptive rotorblade are explained.The understanding of the interaction between the aerodynamic sources and the resulting vibrations and noise is the basis of the design of optimal control concepts.

Compressible potential flow around a helicopter rotor in arbitrary motion

To consider compressible transonic flow effects a vortex lattice method for the computation of helicopter rotor flow is coupled with a field panel method. For this purpose Cartesian grids are used which are not adapted to the contour and only discretise the domain of the nonlinear flow. The basis of this procedure is the separation of the full potential equation into the Laplacian operator and the nonlinear terms. Several methods, including the panel-clustering technique are presented to reduce the computational effort of the method. The developed program ROFPM is validated at several experimental test cases of the Caradonna rotor.

Development of an Experimental Correlation for Transonic Turbine Flow

Optimization of transonic turbine bladings over a broad range of operating conditions calls for better understanding of the relationship between blade profile loss and cascade geometric parameters. In fact, many of the experimental correlations published to date have failed to take into due consideration transonic effects, while others have considered far too few of the numerous geometric parameters affecting profile loss in transonic flows. Through examination of the experimental data gathered by some 20 authors regarding the effects of the most significant blading geometric parameters on profile losses, a loss correlation procedure has been developed. The procedure is especially advantageous in that it allows continuous updating as new experimental data become available.

Measured and calculated airloads on a transport wing model

Wind tunnel measurements of steady and unsteady pressures for a high-aspect-ratio supercritical wing model are compared with calculations by the linear unsteady aerodynamic lifting-surface theory, known as the doublet lattice method, at Mach numbers of 0.60 (subsonic) and 0.78 (transonic). The steady-pressure data comparisons are made for incremental changes in angle of attack and control-surface deflection. The unsteady-pressure data comparisons are made for oscillating control-surface deflections. Some differences between the measured and calculated aerodynamics are attributed to viscous and transonic effects not accounted for in the doublet lattice analysis. Comparisons of the transonic unsteady-pressure data for the oscillating control surfaces are improved by applying empirical corrections, based on the steady-pressure measurements, to the unsteady doublet lattice calculations.

Steady and unsteady transonic airloads on a supercritical wing

A brief description of an experimental investigation to determine steady and unsteady airloads on a fixed and an oscillating supercritical wing model is presented. Steady and unsteady data are correlated with results of computational methods for various representative cases. Steady transonic flow computations were carried out using the XTRAN3S code and the Bailey-Ballhaus/McNally code. Unsteady flow computations were performed using the XTRAN3S code and a quasi-three-dimensional method, which combines the LTRAN2-NLR code with subsonic theory. Compared to the experimental data, the computed results show qualitatively similar transonic effects on pressure distributions and lift coefficients. The accuracy, however, is only modest.

Interaction between a weak oblique shock wave and a turbulent boundary layer in purely supersonic flow

The paper deals with the interaction of a weak oblique shock and a turbulent boundary layer at a flat plate at purely supersonic speeds. Except for a small region where nonlinear transonic effects are of importance analytical expressions for the pressure- and skin friction distributions are derived.

Application of transonic codes to aeroelastic modeling of airfoils including active controls

A study is performed using aeroelastic modeling to investigate the stability behavior of airfoils in small-disturbance transonic flow. Two conventional airfoils, NACA 64.A006 and NACA 64A010, and a supercritical airfoil, MBB A-3, are considered. Three sets of unsteady aerodynamic data are computed using three different transonic codes (LTRAN2-NLR, LTRAN2-HI, and USTS) for comparison purposes. Stability results obtained using a constant matrix, state-space, aeroelastic model are presented in a root-locus format. Use of the state-space model is demonstrated through application to flutter suppression using active controls. Aeroelastic effects due to simple, constant gain, partial feedback, control laws that utilize displacement, velocity, and acceleration sensing are studied using a variety of control gains. Calculations are also performed using linear subsonic aerodynamic theory to reveal the differences between including and not including transonic effects in the aeroelastic model. Aeroelastic stability behavior of these airfoils is physically interpreted and discussed in detail.

Stagnation and transonic effects in thermal blooming

The effects of slewed high-energy beams on an absorbing medium are considered near regions where the relative air velocity is (a) zero and (b) transonic. Limiting effects on the density changes are buoyancy and sound speed changes, respectively. In both cases, the immediate environs of the singular point are greater contributors to phase changes in the beam than are the singularities themselves. Detailed hydrodynamic calculations are carried out for the stagnation region and the steady near-sonic region and compared (in the case of the stagnation region) with a perturbation calculation. Moment methods are used to estimate stagnation-region effects on the subsequent development of the beam.

Aerodynamic Boundary Layer Effects on Flutter and Damping of Plates

References Stable, C. V., Transonic Effects on T-tail Flutter, RM-24, 1959, The Martin Company, Baltimore, Md. Stark, V. J. E., Forces on a Combination of a Wing and a Fin Oscillating in Flow, SAAB TN54, Feb. 1964, Likoping, Sweden. Laschka, B. and Schmid, H., Unsteady Aerodynamic Forces on Coplanar Surfaces in Flow (Wing-Horizontal Tail Interference), AGARD Structures and Materials Panel Meeting, Sept. 25-27,1967, Ottawa, Canada. Topp, L. J., Rowe, W. S., and Shattuck, A. W., Aeroelastic Considerations in the Design of Variable Sweep Airplanes, 1C AS Paper 66-12, Sept. 1966, London, England. Sensberg, O. and Laschka, B., Induced by Aerodynamic between Wing and Tail, Journal of Aircraft, Vol. 7, No. 4, July-Aug. 1970, p. 319-324. Mykytow, W. J., Noll, T. E., Huttsell, L. J., and Shirk, M. H., Investigations Concerning the Coupling Wing-Fuselage-Tail Flutter Phenomenon, Journal of Aircraft, Vol. 9, No. 1, Jan. 1972, pp. 48-54. Albano, E. and Rodden, W. P., A Doublet Lattice Method for Calculating Lift Distributions on Oscillating Surfaces in Flow, AIAA Journal, Vol. 7, No. 2, Feb. 1969, pp. 279-285. Albano, E., Perkinson, F., and Rodden, W. P., Subsonic Surface Theory Aerodynamics and Flutter Analysis of Interfering Wing/Horizontal Tail Configurations, AFFDL-TR-70-59, Sept. 1970, Air Force Flight Dynamics Lab., Wright-Patterson Air Force Base, Ohio. Triplett, W. E., Burkhart, T. H., and Birchfield, E. B., A Comparison of Methods for the Analysis of Wing-Tail Interaction Flutter, AIAA Paper 70-80, 1970, York. Goetz, R. C., Lifting and Control Surface Flutter, TMX52876, Vol. Ill, July 1970, NASA. ^Rodden, W. P., Giesing, J. P., and Kalman, T. P., New Developments and Applications of the Doublet-Lattice Method for Nonplanar Configurations, AGARD Symposium, Tonsberg, Oslofjorden, Norway, Nov. 1970. Chipman, R. R., Interference Flutter on Space Shuttle Configurations, Paper presented at Aerospace Flutter and Dynamics Council Meeting, May 1972, York. Bisplinghoff, R. L. and Ashley, H., Principles of Aeroelasticity, J. Wiley and Sons, York, 1962.

Acceleration of Slender Bodies of Revolution through Sonic Velocity

The linearized theory of slender bodies in arbitrary motion at zero angle of attack has been worked out. The results have been applied to a smooth body accelerating uniformly through sonic velocity. The results theory can be used to estimate the nonlinear or transonic effects. For an accelerating body, the parameter (bl/c2)½ is important where 2b = acceleration, 2l = length of body, c = sound speed at infinity. For sufficiently high (bl/c2)½, transonic effects can be neglected. Using linearized theory to estimate the ratio of nonlinear terms in the differential equation gives λ=nonlinear termssignificant linear terms=34(γ+1)δ2(bl/c2)12{log2δ2(c2bl)12−94},where δ = thickness ratio of body. The result above is evaluated at the maximum thickness of a symmetric parabolic arc body at the instant it passes through sonic velocity. For λ<1 transonic effects can be neglected while for λ>1 they begin to dominate. For practical applications the result shows that there is a possibility of a sufficiently long and slender missile accelerating fast enough to avoid transonic effects (e.g., 50 feet long, 5 percent thick, 3g acceleration). For conventional aircraft, transonic effects will dominate. An interesting side result is that when the acceleration is sufficiently large so that transonic effects do not matter the drag coefficient near sonic speed is independent of the acceleration (CD≐3δ2 for parabolic arc body).