The similitude requirements for the most general test conditions are presented. These similitude requirements are considered in relation to the scaling relationships, test technique, test conditions (including supersonic flow), and test objectives. Particular emphasis is placed on satisfying the various similitude requirements for incompressible and compressible flow conditions. For free flying models tests, the test velocities for incompressible flow are scaled from Froude number similitude requirements and those for compressible flow are scaled from Mach number similitude requirements. The limitations of various test techniques are indicated, with emphasis on the free flying model.

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Similitude requirements and scaling relationships as applied to model testing

The NASA supercritical airfoil development program is summarized in a chronological fashion. Some of the airfoil design guidelines are discussed, and coordinates of a matrix of family related supercritical airfoils ranging from thicknesses of 2 to 18 percent and over a design lift coefficient range from 0 to 1.0 are presented.

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NASA supercritical airfoils: A matrix of family-related airfoils

Some of the advantages and drawbacks of genetic algorithms applications to aerodynamic design are demonstrated. A numerical procedure for the aerodynamic design of transonic airfoils by means of genetic algorithms, with single-point, multipoint, and multiobjective optimization capabilities, is presented. In the ® rst part, an investigation on the relative ef® ciency of different genetic operators combinations is carried out on an aerodynamic inverse design problem. It is shown how an appropriate tuning of the algorithm can provide improved performances, better adaption to design space size and topology, and variables cross correlation. In the second part, the multiobjective approach to design is introduced. The problem of the optimization of the drag rise characteristics of a transonic airfoil is addressed and dealt with using a single point, a multipoint, and a multiobjective approach. A comparison between the results obtained using the three different strategies is ® nally established, showing the advantages of multiobjective optimization.

Inverse and Direct Airfoil Design Using a Multiobjective Genetic Algorithm

Some aspects of aircraft dynamic loads due to flow separation

The test gas for the Langley Pilot Transonic Cryogenic Tunnel is nitrogen. Results from analytical and experimental studies that have verified cryogenic nitrogen as an acceptable test gas are reviewed. Real-gas isentropic and normal-shock flow solutions for nitrogen are compared to the ideal diatomic gas solutions. Experimental data demonstrate that for temperatures above the liquefaction boundaries there are no significant real-gas effects on two-dimensional airfoil pressure distributions. Results of studies to determine the minimum operating temperatures while avoiding appreciable effects due to liquefaction are included.

Cryogenic nitrogen as a transonic wind-tunnel test gas

The theory and advantages of the cryogenic tunnel concept are briefly reviewed. The unique ability to vary temperature independently of pressure and Mach number allows, in addition to large reductions in model loads and tunnel power, the independent determination of Reynolds number, Mach number, and aeroelastic effects on the aerodynamic characteristics of the model. Various combinations of Reynolds number and dynamic pressure can be established to accurately represent flight variations of aeroelastic deformation with altitude changes. The consequences of the thermal and caloric imperfections of the test gas under cryogenic conditions have been examined and found to be insignificant for operating pressures up to 5 atm. The characteristics of the Langley 34 cm (13.5 in.) Pilot Cryogenic Transonic Pressure Tunnel are described and the results of initial tunnel operation are presented. Tests of a two-dimensional airfoil at a Mach number of 0.85 show identical pressure distributions for a chord Reynolds number of 8.6 X 106 obtained first at a stagnation pressure of 4.91 atm at a stagnation temperature of +120°F and then at a stagnation pressure of 1.19 atm at a stagnation temperature of -250°F.

Flight Simulation Characteristics of the Langley High Reynolds Number Cryogenic Transonic Tunnel

The past 6 years of operation with the NASA Langley 0.3 m transonic cryogenic tunnel (TCT) show that there are no insurmountable problems associated with cryogenic testing with gaseous nitrogen at transonic Mach numbers. The fundamentals of the concept were validated both analytically and experimentally and the 0.3 m TCT, with its unique Reynolds number capability, was used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects were developed and cryogenic tunnel conditions can be set and maintained accurately. Cryogenic cooling by injecting liquid nitrogen directly into the tunnel circuit imposes no problems with temperature distribution or dynamic response characteristics. Experience with the 0.3 m TCT, indicates that there is a significant learning process associated with cryogenic, high Reynolds number testing. Many of the questions have already been answered; however, factors such as tunnel control, run logic, economics, instrumentation, and model technology present many new and challenging problems.

Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel

A wind-tunnel research program has been under-taken by the NASA to study the aerodynamic characteristics of T-tail aircraft at high angles of attack. The program was designed to show the effects on longitudinal stability and control of several configuration variables. The results to date do not allow the formulation of general design rules, but the effects of several configuration variables have been noted to have a prime influence on the post-stall characteristics. An increase in tail size, changes in the location of fuselage-mounted engine nacelles, and reduced fuselage-forebody lift were all found to have a beneficial effect on static longitudinal stability at high angles of attack.

Deep-Stall Aerodynamic Characteristics of T-Tail Aircraft

Recent research efforts in the improvement of the maneuverability of fighter aircraft in the high-subsonic and transonic speed range are reviewed with emphasis on the factors affecting aerodynamic boundaries, such as maximum obtainable lift, buffet onset, pitchup, wing rock, and nose slice. The investigations were made using a general research configuration which encompassed a systematic matrix of wing-design parameters. These results illustrated the sensitivity of section and planform geometry to a selected design point. The incorporation of variable-geometry wing devices in the form of flaps or leading-edge slats was shown to provide controlled flow over a wide range of flight conditions and substantial improvements in maneuver capabilities. Additional studies indicated that the blending of a highly swept maneuver strake with an efficient, moderately swept wing offers a promising approach for improving maneuver characteristics at high angles of attack without excessive penalties in structural weight.

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Edward J. Ray

Papers

Cryogenic nitrogen as a transonic wind-tunnel test gas

Flight Simulation Characteristics of the Langley High Reynolds Number Cryogenic Transonic Tunnel

Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel

Deep-Stall Aerodynamic Characteristics of T-Tail Aircraft

Maneuver and buffet characteristics of fighter aircraft