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Emergence of Cybersecurity has increased the focus on security risks to Information Technology (IT) assets going beyond traditional Information Assurance (IA) concerns: More sophisticated threats have emerged from increasing sources as advanced hacker tools and techniques have emerged and proliferated to broaden the attack surface available across globally interconnected networks.
Normal- and oblique-shock flow parameters for air in thermochemical equilibrium are tabulated as a function of shock angle for altitudes ranging from 15.24 km to 91.44 km in increments of 7.62 km at selected hypersonic speeds. Post-shock parameters tabulated include flow-deflection angle, velocity, Mach number, compressibility factor, isentropic exponent, viscosity, Reynolds number, entropy difference, and static pressure, temperature, density, and enthalpy ratios across the shock. A procedure is presented for obtaining oblique-shock flow properties in equilibrium air on surfaces at various angles of attack, sweep, and dihedral by use of the two-dimensional tabulations. Plots of the flow parameters against flow-deflection angle are presented at altitudes of 30.48, 60.96, and 91.44 km for various stream velocities.
This report covers the twelfth of a series of tests conducted to compare different lateral control devices with particular reference to their effectiveness at high angles of attack. The present wind tunnel tests were made with two sizes of upper-surface ailerons on rectangular Clark Y wing models equipped with full span split flaps. The tests showed the effect of the upper-surface ailerons and of the split flaps on the general performance characteristics of the wings, and on the lateral controllability and stability characteristics. The results are compared with those for plain wings with ordinary ailerons of similar sizes.
An engineering-type method is presented for estimating normal-force, axial-force, and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. Static aerodynamic characteristics computed by the method are shown to agree closely with experimental results for slender bodies of circular and elliptic cross section and for winged-circular and winged-elliptic cones. However, the present experimental results used for comparison with the method are limited to angles of attack only up to about 20 deg and Mach numbers from 2 to 4.
Two different singularity methods have been utilized to calculate the potential flow past a three dimensional non-lifting body. Two separate FORTRAN computer programs have been developed to implement these theoretical models, which will in the future allow inclusion of the fuselage effect in a pair of existing subcritical wing design computer programs. The first method uses higher order axial singularity distributions to model axisymmetric bodies of revolution in an either axial or inclined uniform potential flow. Use of inset of the singularity line away from the body for blunt noses, and cosine-type element distributions have been applied to obtain the optimal results. Excellent agreement to five significant figures with the exact solution pressure coefficient value has been found for a series of ellipsoids at different angles of attack. Solutions obtained for other axisymmetric bodies compare well with available experimental data. The second method utilizes distributions of singularities on the body surface, in the form of a discrete vortex lattice. This program is capable of modeling arbitrary three dimensional non-lifting bodies. Much effort has been devoted to finding the optimal method of calculating the tangential velocity on the body surface, extending techniques previously developed by other workers.
As part of a research program aimed at providing information on the high angle-of-attack characteristics of three-surface fighter concepts incorporating a close-coupled canard, an investigation is being conducted on two specific configurations based on the F-18 and F-15 designs. The study configurations are being subjected to a wide range of tests including wind-tunnel tests, dynamic model tests, and piloted simulation. This paper summarizes the results obtained to date in this study. High-alpha results in the areas of static stability, damping, and control characteristics are reviewed and some of the more significant aerodynamic phenomena are identified.
The performance characteristics of a downward canted normal-shock side (scoop) inlet located downstream of a triangular control surface are presented for free-stream Mach numbers of 1.5 and 1.8 in terms of total pressure recovery and mass flow ratio for various boundary-layer removal systems,angles of attack, control surface deflections and adverse yaw. An engine operating condition for a hypothetical turbojet engine is established, and the match point characteristics of the engine-inlet configuration are summarized. 520::It is shown that the diffuser performance increases with increased boundary-layer removal and decreases because of the presence of the wake from the forward control surface. At the higher angles of attack the wake passes over the inlet and does not affect the inlet performance. Adverse yaw reduces the total pressure recovery values below those for the unawed case. Magnitudes of the total pressure recovery were below the theoretical normal-shock recovery for the respective test Mach numbers.
A method is presented for computing normal force and pitching moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. A semiempirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For bodies of revolution, computed aerodynamic characteristics agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9 and for angles of attack from 0 deg to 180 deg. For bodies of elliptic cross section, measured results are predicted well over the investigated Mach number range from 0.6 to 2.0 and the angle range from 0 deg to 60 deg. For all bodies the predictions are best at supersonic Mach numbers. For body-wing and body-wing-tail configurations, measured normal force coefficients and centers are predicted at the upper test Mach number of 2.0. As the Mach number is decreased to 0.6, the agreement for the normal-force coefficients rapidly deteriorates. When model flow-separation and vortex patterns are asymmetric, undesirable side forces are usually measured on the models at subsonic Mach numbers and zero sideslip angle. Generally, the side-force coefficients decrease or vanish with: increase in Mach number, decrease in nose fineness ratio, nose blunting, and flattening of body cross section.
A wind tunnel investigation was conducted to study the low speed, high angle of attack stability characteristics of a three surface fighter concept based on the F-15 configuration. Static force data were measured over angle of attack and side-slip ranges of 0 to 85 and -10 and 10 deg, respectively. A force oscillation technique was used to obtain dynamic derivatives at angles of attack from 0 to 60 deg. The tests were conducted for several canard deflections and with the canards removed to investigate the effects of the close coupled canard on the high angle of attack stability characteristics of the configuration. A fuselage strake was developed which significantly improved static lateral directional stability characteristics at high angles of attack while also increasing the maximum lift of the configuration.
An approximate solution technique has been developed for three-dimensional, inviscid, hypersonic flows. The method uses Maslen's explicit pressure equation and the assumption of approximate stream surfaces in the shock layer. This approximation represents a simplification of Maslen's asymmetric method. The solution procedure involves iteratively changing the shock shape in the subsonic-transonic region until the correct body shape is obtained. Beyond this region, the shock surface is determined by using a marching procedure. Results are presented herein for a paraboloid and elliptic cone at angle of attack. Calculated surface pressure distributions, shock shapes, and property profiles are compared with experimental data and finite-difference solutions of the Euler equations. Comparisons of the results of the present method with experimental data and detailed predictions are very good. Since the present method provides a very rapid computational procedure, it can be used for parametric or preliminary design applications. One useful application would be to incorporate a heating procedure for aerothermal studies.
A 10% scale, high-lift version of the Common Research Model (CRM-HL) was tested in the NASA Langley Research Center 14-by22-Foot Subsonic Tunnel. The focus of the wind tunnel test campaign was to assess the feasibility of active flow control (AFC) on the simplified CRM-HL configuration. The modular design of the CRM-HL model enabled the testing of the conventional CRM-HL version to establish a benchmark for the AFC research. The wind tunnel model has the capability of being tested with and without the engine nacelle/pylon for comparison. The wind tunnel investigation included the acquisition of force and moment data, steady pressure data, as well as surface flow visualization using minitufts. For some select cases, numerical simulations were performed to aid in understanding the wind tunnel data. Wind tunnel measurements indicated a lift degradation of the conventional CRM-HL at high angles of attack. The surface flow visualization with fluorescent minitufts revealed that the lift degradation is due to flow separation caused by the nacelle/pylon wake on the inboard wing. The wake and the consequent flow separation were successfully mitigated using a nacelle chine installed on the inboard side of the engine nacelle.
Computational analysis of flow over the F/A-18 aircraft is presented along with complementary data from both flight and wind tunnel experiments. The computational results are based on the three-dimensional thin-layer Navier-Stokes formulation and are obtained from an accurate surface representation of the fuselage, leading-edge extension (LEX), and the wing geometry. However, the constraints imposed by either the flow solver and/or the complexity associated with the flow-field grid generation required certain geometrical approximations to be implemented in the present numerical model. In particular, such constraints inspired the removal of the empennage and the blocking (fairing) of the inlet face. The results are computed for three different free-stream flow conditions and compared with flight test data of surface pressure coefficients, surface tuft flow, and off-surface vortical flow characteristics that included breakdown phenomena. Excellent surface pressure coefficient correlations, both in terms of magnitude and overall trend, are obtained on the forebody throughout the range of flow conditions. Reasonable pressure agreement was obtained over the LEX; the general correlation tends to improve at higher angles of attack. The surface tuft flow and the off-surface vortex flow structures compared qualitatively well with the flight test results. To evaluate the computational results, a wind tunnel investigation was conducted to determine the effects of existing configurational differences between the flight vehicle and the numerical model on aerodynamic characteristics. In most cases, the geometrical approximations made to the numerical model had very little effect on overall aerodynamic characteristics.
As part of NATO STO AVT-346 “Predicting Hypersonic Boundary-Layer Transition on Complex Geometries,” coordinated experimental and computational studies were conducted on the ROTEX-T, a cone-flare geometry used in a successful flight-test experiment. At the as-flown conditions, a separation bubble existed at the compression corner. Separation, reattachment, and the multifaceted linear instability paths leading this bubble to transition to turbulence are challenging to predict, but have significant impact on surface pressure and heat flux. High-resolution background-oriented schlieren and infrared thermography measurements were made in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel at freestream unit Reynolds numbers from 5.8×10(exp 6) to 12.2×10(exp 6) m-1and nominally zero angle of attack. High-speed self-aligned focusing schlieren, infrared thermography, and focused laser differential interferometry measurements were made in the AFRL Mach-6 Ludwieg Tube from 2.2×10(exp 6) to 24.7×10(exp 6) m-1and nominally zero degrees angle of attack. The surface heat-flux and Stanton number distributions were computed. Separation and reattachment locations, as well as the flow state at each, were determined from the combination of surface and off-wall measurements. The convective and global boundary-layer instabilities of the axisymmetric laminar flow at the experimental conditions were investigated computationally. Amplification of Mack’s first and second modes were observed to have logarithmic amplification factors between 5 to 7.5 at the separation location, depending on conditions. The flow was found to be globally unstable to stationary three-dimensional disturbances concentrated in the reattachment region. Previous analysis of the ROTEX-T flight data had not assessed reattachment location or the flow state upon reattachment. Thanks to the insights gained from the coordinated, on- and off-wall ground-test measurements, these evaluations have now been made. The separation location indicated by laminar simulations is consistently numerically predicted to be upstream of the experimentally observed location for a transitional separation bubble. The cause of this difference is understood to lie in the steady-state and axisymmetric assumptions made by both solvers employed to compute the basic states analyzed, as flow topology considerations assert that unsteady two-dimensional or axisymmetric separation bubbles are structurally unstable and will become three-dimensional. Computed laminar heating rates prior to separation agreed well with experiment; transitional heating rates after reattachment were between laminar and turbulent computations.
As part of NATO STO AVT-346 “Predicting Hypersonic Boundary-Layer Transition on Complex Geometries,” coordinated experimental and computational studies were conducted on the ROTEX-T, a cone-flare geometry used in a successful flight-test experiment. At the as-flown conditions, a separation bubble existed at the compression corner. Separation, reattachment, and the multifaceted linear instability paths leading this bubble to transition to turbulence are challenging to predict, but have significant impact on surface pressure and heat flux. High-resolution background-oriented schlieren and infrared thermography measurements were made in the AFOSR–Notre Dame Large Mach-6 Quiet Tunnel at freestream unit Reynolds numbers from 5.8×10(exp 6) to 12.2×10(exp 6) m-1and nominally zero angle of attack. High-speed self-aligned focusing schlieren, infrared thermography, and focused laser differential interferometry measurements were made in the AFRL Mach-6 Ludwieg Tube from 2.2×10(exp 6) to 24.7×10(exp 6) m-1and nominally zero degrees angle of attack. The surface heat-flux and Stanton number distributions were computed. Separation and reattachment locations, as well as the flow state at each, were determined from the combination of surface and off-wall measurements. The convective and global boundary-layer instabilities of the axisymmetric laminar flow at the experimental conditions were investigated computationally. Amplification of Mack’s first and second modes were observed to have logarithmic amplification factors between 5 to 7.5 at the separation location, depending on conditions. The flow was found to be globally unstable to stationary three-dimensional disturbances concentrated in the reattachment region. Previous analysis of the ROTEX-T flight data had not assessed reattachment location or the flow state upon reattachment. Thanks to the insights gained from the coordinated, on- and off-wall ground-test measurements, these evaluations have now been made. The separation location indicated by laminar simulations is consistently numerically predicted to be upstream of the experimentally observed location for a transitional separation bubble. The cause of this difference is understood to lie in the steady-state and axisymmetric assumptions made by both solvers employed to compute the basic states analyzed, as flow topology considerations assert that unsteady two-dimensional or axisymmetric separation bubbles are structurally unstable and will become three-dimensional. Computed laminar heating rates prior to separation agreed well with experiment; transitional heating rates after reattachment were between laminar and turbulent computations.
A wind tunnel experiment was performed on a delta wing with a leading edge sweep of 70 deg. Unsteady pressures were measured on the suction surface of the wing as it was oscillated from 0-30 deg, and 2-60 deg angle-of-attack. Pressure coefficients were measured at different surface locations for two pitch rates. Static pressure measurements were also obtained for comparison with the dynamic results. The Reynolds number was fixed at 420,000, based on the centerline chord length. Pressure measurements were made from 35-90 percent of the chord, along a ray from the apex at 60 percent of the local semispan. Spanwise measurements were also made at a constant chord location, x/c = 75 percent from the apex. The unsteady pressure data over the 0-30 deg angle-of-attack range showed pressures fluctuating in phase with model motion, and little overshoot from the static values. The pressures for the large amplitude motion showed large overshoots from the static values. In addition, during the high angle-of-attack portion of the motion, the upstroke (angle of attack increasing) pressure coefficients were typically much lower than the downstroke values. For the lower pitch rate, there was little difference between upstroke and downstroke pressures at the low angles of attack.
The Magicote C process developed by S.M. Thompsen was evaluated for use in applying an antireflective coating to the cover plates of solar panels. The process uses a fluosilicic acid solution supersaturated with silica at elevated temperature to selectively attack the surface of soda-lime glass cover plates and alter the physical and chemical composition of a thin layer of glass. The altered glass layer constitutes an antireflective coating. The process produces coatings of excellent optical quality which possess outstanding resistance to soiling and staining. The coatings produced are not resistant to mechanical abrasion and are attacked to some extent by glass cleansers. Control of the filming process was found to be difficult.
Noise spectra were measured in three dimensions for several surfaces immersed in turbulent flow from a jet and over a range of flow conditions. The data are free field and were corrected to remove the small contributions of jet noise, atmospheric attenuation and feedback tones. These broadband data were compared with the results of available theories which are only strictly applicable to simple geometries over a limited range of conditions. The available theories proved to be accurate over the range of flow, chord length, thickness, angle of attack, and surface geometries defined by the experiments. These results apply to the noise generated by fixed surfaces in engine passages, the lifting surfaces of aircraft and also to fan noise.
Noise spectra were measured in three dimensions for several surfaces immersed in turbulent flow from a jet and over a range of flow conditions. The data are free field and were corrected to remove the small contributions of jet noise, atmospheric attenuation and feedback tones. These broadband data were compared with the results of available theories which are only strictly applicable to simple geometries over a limited range of conditions. The available theories proved to be accurate over the range of flow, chord length, thickness, angle of attack, and surface geometries defined by the experiments. These results apply to the noise generated by fixed surfaces in engine passages, the lifting surfaces of aircraft and also to fan noise.