ALBERT

Description: This investigation concerns the time and frequency formulations of non-linear two-dimensional lifting surfaces exposed to an incompressible flow field and subjected to an external pressure pulse. In order to address this problem, Volterra series approach in conjunction with the multidimensional Laplace transform is used. This methodology enabling one to solve the aeroelastic governing equations of lifting surfaces opens the way to connect this methodology with that based on neural networks and NARMAX/NARX networks models. Moreover, this extended way to address this problem constitutes a good basis for treatment of the theory of 3D lifting surfaces.

Description: The use of balloons/aerobots on Mars has been under consideration for many years. Concepts include deployment during entry into the atmosphere from a carrier spacecraft, deployment from a lander, use of super-pressurized systems for long duration flights, 'hot-air' systems, etc. Principal advantages include the ability to obtain high-resolution data of the surface because balloons provide a low-altitude platform which moves relatively slowly. Work conducted within the last few years has removed many of the technical difficulties encountered in deployment and operation of balloons/aerobots on Mars. The concept proposed here (a tethered balloon released from a lander) uses a relatively simple approach which would enable aspects of Martian balloons to be tested while providing useful and potentially unique science results. Tethered Micro-Balloons on Mars (TMBM) would be carried to Mars on board a future lander as a stand-alone experiment having a total mass of one to two kilograms. It would consist of a helium balloon of up to 50 cubic meters that is inflated after landing and initially tethered to the lander. Its primary instrumentation would be a camera that would be carried to an altitude of up to tens of meters above the surface. Imaging data would be transmitted to the lander for inclusion in the mission data stream. The tether would be released in stages allowing different resolutions and coverage. In addition during this staged release a lander camera system may observe the motion of the balloon at various heights above he lander. Under some scenarios upon completion of the primary phase of TMBM operations, the tether would be cut, allowing TMBM to drift away from the landing site, during which images would be taken along the ground.

Description: In the early 1990's, Glezer and his co-workers at Georgia Tech made a startling discovery. They found that forcing at frequencies too high to directly affect the production scales led to a dramatic alteration in the development of a turbulent shear layer. An experimental study of this phenomenon is presented in Wiltse and Glezer. They used piezoelectric actuators located near the jet exit plane to force the shear layers of a square low-speed jet. The actuators were driven at a high frequency in the Kolmogorov inertial subrange, much higher than the frequencies associated with the large-scale motion (where the turbulent energy is produced and located) but much lower than those associated with the Kolmogorov scale (where the turbulent energy is dissipated). Measurements of the shear-layer turbulence showed that direct excitation of small-scale motion by high-frequency forcing led to an increase in the turbulent dissipation of more than an order of magnitude in the initial region of the shear layer! The turbulent dissipation gradually decreased with downstream distance but remained above the corresponding level for the unforced flow at all locations examined. The high-frequency forcing increased the turbulent kinetic energy in the initial region near the actuators, but the kinetic energy decreased quite rapidly with downstream distance, dropping to levels that were a small fraction of the level for the unforced case. Perhaps most importantly from the present standpoint, the high-frequency forcing significantly decreased the energy in the large-scale motion, increasingly so with downstream distance. Wiltse and Glezer interpreted this behavior as an enhanced transfer of energy from the large scales to the small scales. The initial work by Wiltse and Glezer has expanded into other applications. To explore the potential of high-frequency forcing for active acoustic suppression, in 1998 the first author proposed a set of experiments involving an edge tone shear layer and an open cavity flow. This work was funded by the US Air Force Research Laboratory, and the experiments were developed and executed at Boeing by Raman and Kibens. These experiments involved high-frequency forcing applied to low-speed flows using wedge piezo actuators and powered resonance tubes. The system is simple, open loop, compact, potentially requires little power, and is easily integrated. Dramatic results, such as reductions of 20 dB in spectral peaks and 5-8 dB in overall levels across the entire acoustic spectrum, were obtained in some cases. Sample results are presented. Following this success in low-speed flows, an international cooperative program continuing this work involved transonic experiments in a mid-size facility in the United Kingdom. Similar reductions in noise level were obtained in these transonic experiments. Discussion of this work is given in Raman et at. and Stanek, Raman, Kibens, and Ross. Other experiments at Georgia Tech have shown significant potential of high-frequency forcing in controlling reaction rates in chemically reacting flows.

Description: NASA researchers have designed a system to predict aircraft wake turbulence on final approach, so airliners can be spaced more safely and efficiently. This technology, known as the Aircraft VOrtex Spacing System (AVOSS), demonstrates an integration of technologies that provides weather-dependent dynamic aircraft spacing for wake avoidance in a real-time relevant environment. AVOSS was successfully demonstrated at Dallas Fort-Worth Airport in July 2000. The demonstration represented the culmination of 6 years of field-testing, data collection, and development.

Description: The aerodynamic development of an engine inlet requires a comprehensive program of both wind tunnel testing and Computational Fluid Dynamics (CFD) simulations. To save time and resources, much "testing" is done using CFD before any design ever enters a wind tunnel. The focus of my project this summer is on CFD analysis tool development. In particular, I am working to further develop the capabilities of the Planar Inlet Design and Analysis Process (PINDAP). "PINDAP" is a collection of computational tools that allow for efficient and accurate design and analysis of the aerodynamics about and through inlets that can make use of a planar (two-dimensional or axisymmetric) geometric and flow assumption. PINDAP utilizes the WIND CFD flow solver, which is capable of simulating the turbulent, compressible flow field. My project this summer is a continuation of work that I performed for two previous summers. Two years ago, I used basic features of the PINDAP to design a Mach 5 hypersonic scramjet engine inlet and to demonstrate the feasibility of the PINDAP. The following summer, I worked to develop its geometry and grid generation capabilities to include subsonic and supersonic inlets, complete bodies and cowls, conic leading and trailing edges, as well as airfoils. These additions allowed for much more design flexibility when using the program.

Description: Aircraft travel has become a major form of transportation. Several of our major airports are operating near their capacity limit, increasing congestion and delays for travelers. As a result, the National Aeronautics and Space Administration (NASA) has been working in conjunction with the Federal Aviation Administration (FAA), airline operators, and the airline industry to increase airport capacity without sacrificing public safety. One solution to the problem is to increase the number of airports and build new. runways; yet, this solution is becoming increasingly difficult due to limited space. A better solution is to increase the production per runway. This solution increases the possibility that one aircraft will encounter the trailing wake of another aircraft. Hazardous wake vortex encounters occur when an aircraft encounters the wake produced by a heavier aircraft. This heavy-load aircraft produces high-intensity wake turbulence that redistributes the aerodynamic loads of trailing smaller aircraft. This situation is particularly hazardous for smaller aircraft during takeoffs and landings. In order to gain a better understanding of the wake-vortex/aircraft encounter phenomena, NASA Langley Research Center conducted a series of flight tests from 1995 through 1997. These tests were designed to gather data for the development a wake encounter and wake-measurement data set with the accompanying atmospheric state information. This data set is being compiled into a database that can be used by wake vortex researchers to compare with experimental and computational results. The purpose of this research is to derive and implement a procedure for calculating the wake-vortex/aircraft interaction portion of that database by using the data recorded during those flight tests. There were three objectives to this research. Initially, the wake-induced forces and moments from each flight were analyzed based on varying flap deflection angles. The flap setting alternated between 15 and 30 degrees while the separation distance remained constant. This examination was performed to determine if increases in flap deflection would increase or decrease the effects of the wake-induced forces and moments. Next, the wake-induced forces and moments from each flight were analyzed based on separation distances of 1-3 nautical miles. In this comparison, flap deflection was held constant at 30 degrees. The purpose of this study was to determine if increased separation distances reduced the effects of the wake vortex on the aircraft. The last objective compared the wake-induced forces and moments of each flight as it executed a series of maneuvers through the wake-vortex. This analysis was conducted to examine the impact of the wake on the B737 as it traversed the wake horizontally and vertically. Results from the first analysis indicated that there was no difference in wake effect at flap deflections of 15 and 30 degrees. This conclusion is evidenced in the cases of the wake-induced sideforce, rolling moment, and yawing moment. The wake-induced lift, drag, and pitching moment cases yielded less conclusive results. The second analysis compared the wake-induced forces and moments at separation distances of 1-3 nautical miles. Results indicated that there was no significant difference in the wake-induced lift, drag, sideforce, or yawing moment coefficients. The analysis compared the wake-induced forces and moments based on different flight maneuvers. It was found that the wake-induced forces and moments had the greatest impact on out-to-in and in-to-out maneuvers.

Description: Effect of rapid pressure decay on solid propellant combustion

Description: Several improvements have recently been made in the thermal analysis methods for leading edges of a hypersonic vehicle. The leading edges of this vehicle undergo exceptionally high heat loads that incorporate extreme spatial gradients as well as severe transients. Due to the varying flight conditions, complex geometry, and need for thermal loads at many points along the trajectory, full computational fluid dynamics (CFD) analysis of the aeroheating; loads is not feasible. Thus, engineering methods must be used to determine the aeroheating on the vehicle surfaces, and that must be utilized in the thermal analysis. Over the last year, the thermal analysis of a hypersonic vehicle has been enhanced in several ways. Two different engineering codes are used to predict aeroheating loads: one over the curve near the stagnation point, and the other on flat surfaces downstream of the leading edge. These two are matched together at the intersection point using a method that allows closer approximation of CFD results. User-developed FORTRAN, which is part of the thermal solver PATRAN Thermal, is used to accomplish this. The customizable FORTRAN code also allows use of many different time- and space-dependent factors, interpolation of the heat load in time and space, and inclusion of both highly swept and unswept grid structures. This FORTRAN is available to other PATRAN users who may want to accomplish a similar objective in analysis. Flux, rather than convective coefficient, is used to define heat loads, which allows more accurate analysis as well as better application of margins. Improvements have also been made in more efficient utilization of imported CAD geometry, by creating faces on solids to facilitate load application.

Description: The Aerothermodynamics Branch at NASA - Langley Research Center is tasked with developing, assessing and applying aerothermodynamic technologies to enable the development of hypersonic aircraft, launch vehicles, and planetary/earth entry systems. To accomplish this mission, the Branch capitalizes on the synergism between the experimental and computational facilities/tools which reside in the branch and a staff that can draw on five decades of experience in aerothermodynamics. The Aerothermodynamics Branch is staffed by 30 scientists/engineers. The staff, of which two-thirds are less than 40 years old, is split evenly between experimentalists and computationalists. Approximately 90 percent of the staff work on space transportation systems while the remainder work on planetary missions. The Branch manages 5 hypersonic wind tunnels which are staffed by 14 technicians, numerous high end work stations and a SGI Origin 2000 system. The Branch also utilizes other test facilities located at Langley as well as other national and international test sites. Large scale computational requirements are met by access to Agency resources.

Description: The purpose of health monitoring systems is to detect failures or defects for increased safety and performance and to provide on-condition maintenance with reduced costs. The problems associated with health monitoring systems include high rates of false alarms and missed failures, which make monitoring an unreliable and costly task. The reason for this is that unaccounted variations invalidate signal modeling assumptions. Our approach was to focus on vibration monitoring of rotating components. We analyzed baseline signals to determine statistical variations, identify and model factors that influence vibrations (pre-production vs. post-production variations), determine hit and false alarm rates with baseline flight data, model and predict effects of defects and variations on vibrations, and develop algorithms and metrics for failure and anomaly detection in the presence of variations.

Description: Two constrained shape optimization problems relevant to aerodynamics are solved by genetic programming, in which a population of computer programs evolves automatically under pressure of fitness-driven reproduction and genetic crossover. Known optimal solutions are recovered using a small, naive set of elementary operations. Effectiveness is improved through use of automatically defined functions, especially when one of them is capable of a variable number of iterations, even though the test problems lack obvious exploitable regularities. An attempt at evolving new elementary operations was only partially successful.

Description: A method for real-time estimation of parameters in a linear dynamic state-space model was developed and studied. The application is aircraft dynamic model parameter estimation from measured data in flight. Equation error in the frequency domain was used with a recursive Fourier transform for the real-time data analysis. Linear and nonlinear simulation examples and flight test data from the F-18 High Alpha Research Vehicle were used to demonstrate that the technique produces accurate model parameter estimates with appropriate error bounds. Parameter estimates converged in less than one cycle of the dominant dynamic mode, using no a priori information, with control surface inputs measured in flight during ordinary piloted maneuvers. The real-time parameter estimation method has low computational requirements and could be implemented

Description: This presentation gives information on: pitch criteria based on airplane Bandwidth; apply research, experimental, operational data; compare Smith-Geddes, Gibson, Neal-Smith criteria; bandwidth criteria for Category II PIO; control/response sensitivity and PIO; extension to roll a axis; and some recommendations.

Description: This presentation discuss ground-based simulation, and flight-tests study (HAVE PIO). The purpose of this study is to determine if the amount of platform motion affects ability to replicate in-flight results.

Description: Buffeting is an aeroelastic phenomenon which plagues high performance aircraft, especially those with twin vertical tails like the F/A-18, at high angles of attack. This buffeting is a concern from fatigue and inspection points of view. By means of wind-tunnel and flight tests, this phenomenon is well studied to the point that buffet loads can be estimated and fatigue life can be increased by structural enhancements to the airframe. In more recent years, buffeting alleviation through active control of smart materials has been highly researched in wind-tunnel proof-of-concept demonstrations and full-scale ground tests using the F/A-18 as a test bed. Because the F/A-18 resides in fleets outside as well as inside the United States, these tests have evolved into international collaborative research activities with Australia and Canada, coordinated by the Air Force Research Laboratory (AFRL) and conducted under the auspices of The Technical Cooperation Program (TTCP). With the recent successes and advances in smart materials, the main focus of these buffeting alleviation tests has also evolved to a new level: utilize the F/A-18 as a prototype to mature smart materials for suppressing vibrations of aerospace structures. The role of the NASA Langley Research Center (LaRC) in these programs is presented.

Description: This viewgraph presentation provides information on the status of nozzle aerodynamic technology at MSFC (Marshall Space Flight Center). The objectives of this presentation were to provide insight into MSFC in-house nozzle aerodynamic technology, design, analysis, and testing. Under CDDF (Center Director's Discretionary Fund), 'Altitude Compensating Nozzle Technology', are the following tasks: Development of in-house ACN (Altitude Compensating Nozzle) aerodynamic design capability; Building in-house experience for all aspects of ACN via End-to-End Nozzle Test Program; Obtaining Experimental Data for Annular Aerospike: Thrust eta, TVC (thrust vector control) capability and surface pressures. To support selection/optimization of future Launch Vehicle propulsion we needed a parametric design and performance tool for ACN. We chose to start with the ACN Aerospike Nozzles.

Description: The Aerodynamic Research Facility (ARF) LGBB (Liquid Glide-back Booster) Stage Separation Test is part of the Multi-Center Second Generation In-House Tool Development Task. The ARF LGBB Stage Separation Test has been completed at MSFC (Marshall Space Flight Center). It includes the following: PSP (Project Study Plan) Feasibility Test; Isolated Force/Moment Data; Bimese Configuration Force/Moment Data; Schlieren Video. The LGBB Bimese Reference Configuration Analyses and Test Results In-Work to Develop Tools and Database. Preliminary results showed qualitative agreement with CFD (computational fluid dynamics) aerodynamic predictions. The preliminary results exhibit the complex nature of the stage separation aerothermal problem.

Description: The effect of mass addition on the flow over bodies moving at hypersonic speeds has been studied by several investigators (e.g., Cresci and Libby). In most of this work, primary attention logically has been directed toward the effects of foreign-gas injection on heat transfer and pressure distributions, and, principally for this reason, most of the work ha been done at zero angle of attack. The foreign gas can be provided either by some active injection system or by the action of an ablation heat shield. With increasing rates of injection, the basic flow about the body can be affected significantly. One such effect was observed in the paper by Cresci and Libby, where it was shown that the shockwave standoff distance can be increased by gas injection at the nose of a body.

Description: The paper summarizes the state of the art in aeronautical drag reduction across the speed range for the conventional drag components of viscous drag, drag due to lift and wave drag. It also describes several emerging drag-reduction approaches that are either active or reactive/interactive and the drag reduction potentially available from synergistic combinations of advanced configuration aerodynamics, viscous drag-reduction approaches, revolutionary structural concepts and propulsion integration.

Description: In an effort to discover the causes for disagreement between previous two-dimensional (2-D) computations and nominally 2-D experiment for flow over the three-element McDonnell Douglas 30P-30N airfoil configuration at high lift, a combined experimental/CFD investigation is described. The experiment explores several different side-wall boundary layer control venting patterns, documents venting mass flow rates, and looks at corner surface flow patterns. The experimental angle of attack at maximum lift is found to be sensitive to the side-wall venting pattern: a particular pattern increases the angle of attack at maximum lift by at least 2 deg. A significant amount of spanwise pressure variation is present at angles of attack near maximum lift. A CFD study using three-dimensional (3-D) structured-grid computations, which includes the modeling of side-wall venting, is employed to investigate 3-D effects on the flow. Side-wall suction strength is found to affect the angle at which maximum lift is predicted. Maximum lift in the CFD is shown to be limited by the growth of an off-body corner flow vortex and consequent increase in spanwise pressure variation and decrease in circulation. The 3-D computations with and without wall venting predict similar trends to experiment at low angles of attack, but either stall too early or else overpredict lift levels near maximum lift by as much as 5%. Unstructured-grid computations demonstrate that mounting brackets lower the lift levels near maximum lift conditions.

Description: As a portion of the Benchmark Models Program at NASA Langley, a simple generic model was developed for active controls research and was called BACT for Benchmark Active Controls Technology model. This model was based on the previously-tested Benchmark Models rectangular wing with the NACA 0012 airfoil section that was mounted on the Pitch and Plunge Apparatus (PAPA) for flutter testing. The BACT model had an upper surface spoiler, a lower surface spoiler, and a trailing edge control surface for use in flutter suppression and dynamic response excitation. Previous experience with flutter suppression indicated a need for measured control surface aerodynamics for accurate control law design. Three different types of flutter instability boundaries had also been determined for the NACA 0012/PAPA model, a classical flutter boundary, a transonic stall flutter boundary at angle of attack, and a plunge instability near M = 0.9. Therefore an extensive set of steady and control surface oscillation data was generated spanning the range of the three types of instabilities. This information was subsequently used to design control laws to suppress each flutter instability. There have been three tests of the BACT model. The objective of the first test, TDT Test 485, was to generate a data set of steady and unsteady control surface effectiveness data, and to determine the open loop dynamic characteristics of the control systems including the actuators. Unsteady pressures, loads, and transfer functions were measured. The other two tests, TDT Test 502 and TDT Test 5 18, were primarily oriented towards active controls research, but some data supplementary to the first test were obtained. Dynamic response of the flexible system to control surface excitation and open loop flutter characteristics were determined during Test 502. Loads were not measured during the last two tests. During these tests, a database of over 3000 data sets was obtained. A reasonably extensive subset of the data sets from the first two tests have been chosen for Test Cases for computational comparisons concentrating on static conditions and cases with harmonically oscillating control surfaces. Several flutter Test Cases from both tests have also been included. Some aerodynamic comparisons with the BACT data have been made using computational fluid dynamics codes at the Navier-Stokes level (and in the accompanying chapter SC). Some mechanical and active control studies have been presented. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. Cases for static angles of attack, static trailing-edge and upper-surface spoiler deflections are included for a range of conditions near those for the oscillation cases. Cases for trailing-edge control and upper-surface spoiler oscillations for a range of Mach numbers, angle of attack, and static control deflections are included. Cases for all three types of flutter instability are selected. In addition some cases are included for dynamic response measurements during forced oscillations of the controls on the flexible mount. An overview of the model and tests is given, and the standard formulary for these data is listed. Some sample data and sample results of calculations are presented. Only the static pressures and the first harmonic real and imaginary parts of the pressures are included in the data for the Test Cases, but digitized time histories have been archived. The data for the Test Cases are also available as separate electronic files.

Description: The Benchmark Active Controls Technology (BACT) wing test (see chapter 8E) provides data for the validation of aerodynamic, aeroelastic, and active aeroelastic control simulation codes. These data provide a rich database for development and validation of computational aeroelastic and aeroservoelastic methods. In this vein, high-level viscous CFD analyses of the BACT wing have been performed for a subset of the test conditions available in the dataset. The computations presented in this section investigate the aerodynamic characteristics of the rigid clean wing configuration as well as simulations of the wing with a static and oscillating aileron and spoiler deflection. Two computational aeroelasticity codes extensively used at NASA Langley Research Center are implemented in this simulation. They are the ENS3DAE and CFL3DAE computational aeroelasticity programs. Both of these methods solve the three-dimensional compressible Navier-Stokes equations for both rigid and flexible vehicles, but they use significantly different approaches to the solution 6f the aerodynamic equations of motion. Detailed descriptions of both methods are presented in the following section.

Description: Circulation control wings are a type of pneumatic high-lift device that have been extensively researched as to their aerodynamic benefits. However, there has been little research into the possible airframe noise reduction benefits of a circulation control wing. The key element of noise is the jet noise associated with the jet sheet emitted from the blowing slot. This jet sheet is essentially a high aspect-ratio rectangular jet. A recent study on high aspect-ratio jet noise was performed on a nozzle with aspect-ratios ranging from 100 to 3,000. In addition to the acoustic data, fluid dynamic measurements were made as well. This paper uses the results of these two studies and attempts to develop a prediction scheme for high aspect-ratio jet noise

Description: The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.

Description: This portion of the report documents the results of an experimental program, which focused on pulsed blowing from the trailing edge of a CCW. The main objective of this study was to assess whether pulsed blowing resulted in more, less, or the same amount of radiated noise to the farfield. Results show that a reduction in far-field noise of up to 5 dB is measured when pulse flow is compared to steady flow for an equivalent lift configuration. This reduction is in the spectral region associated with the trailing edge jet noise. This result is due to the unique advantage that pulsed flow has over steady flow. For a range of frequencies, more lift is experienced with the same mass flow as the steady case. Thus, for an equivalent lift and slot height, the pulsed system can operate at lower jet velocities, and hence lower jet noise. At low frequencies (below 1 kHz), the pulsed flow configuration generated more noise in the farfield. This is most likely due to the pulsing mechanism itself. Since the high pressure air feeding the pulsing mechanism was first passed through a high performance muffler, it is likely that this increase in not due to upstream valve noise. Most likely, the impulsive component of the air that periodically fills the plenum causes a broadband source that reaches the farfield. Although the benefit of a pulse trailing edge jet is evident from a mass flow usage and jet noise perspective, attention should be paid towards the design of a viable pulsing system. Future research program in this area should concentrate on the development of a "quiet" pulsing device.

Description: Circulation control wings are a type of pneumatic high-lift device that have been extensively researched as to their aerodynamic benefits. However, there has been little research into the possible airframe noise reduction benefits of a circulation control wing. The key element of noise is the jet noise associated with the jet sheet emitted from the blowing slot. High aspect-ratio jet acoustic results (aspect-ratios from 100 to 3,000) from a related study showed that the jet noise of this type of jet was proportional to the slot height to the 3/2 power and slot width to the 1/2 power. Fluid dynamic experiments were performed in the present study on the high aspect-ratio nozzle to gain understanding of the flow characteristics in an effort to relate the acoustic results to flow parameters. Single hot-wire experiments indicated that the jet exhaust from the high aspect-ratio nozzle was similar to a 2-d turbulent jet. Two-wire space-correlation measurements were performed to attempt to find a relationship between the slot height of the jet and the length-scale of the flow noise generating turbulence structure. The turbulent eddy convection velocity was also calculated, and was found to vary with the local centerline velocity, and also as a function of the frequency of the eddy.

Description: This Appendix documents the salient results from an effort to mitigate the so-called flap-edge noise generated at the split between a flap edge that is deployed and the undeployed flap. Utilizing a Coanda surface installed at the flap edge, steady blowing was used in an attempt to diminish the vortex strength resulting from the uneven lift distribution. The strength of this lifting vortex was augmented by steady blowing over the deployed flap. The test article for this study was the same 2D airfoil used in the steady blowing program reported earlier (also used in pulsed blowing tests, see Appendix G), however its trailing edge geometry was modified. An exact duplicate of the airfoil shape was made out of fiberglass with no flap, and in the clean configuration. It was attached to the existing airfoil to make an airfoil that has half of its flap deployed and half un-deployed. Figure 1 shows a schematic of the planform showing the two areas where steady blowing was introduced. The flap-edge blowing or the auxiliary blowing was in the direction normal to the freestream velocity vector. Slot heights for the blowing chambers were on the order of 0.0 14 inches.

Description: We have designed and built an instrument to measure and monitor the "nightglow" of the Earth's atmosphere in the near ultraviolet (NUV). In this paper we describe the design of this instrument, called NIGHTGLOW. NIGHTGLOW is designed to be flown-from a high altitude research balloon, and circumnavigate the globe. NIGHTGLOW is a NASA, University of Utah, and New Mexico State University project. A test flight took place from Palestine, Texas on July 5, 2000, lasting about 8 hours. The instrument performed well and landed safely in Stiles, Texas with little damage. The resulting measurements of the NUV nightglow are consistent with previous measurements from sounding rockets and balloons. The results will be presented and discussed.

Description: Motivation - Higher loading on Low-Pressure Turbine (LPT) airfoils: Reduce airfoil count, weight, cost. Increase efficiency, and Limited by suction side separation. Growing understanding of transition, separation, wake effects: Improved models. Take advantage of wakes. Higher lift airfoils in use. Further loading increases may require flow control: Passive: trips, dimples, etc. Active: plasma actuators, vortex generator jets (VGJs). Can increased loading offset higher losses on high lift airfoils. Objectives: Advance knowledge of boundary layer separation and transition under LPT conditions. Demonstrate, improve understanding of separation control with pulsed VGJs. Produce detailed experimental data base. Test and develop computational models.

Description: In assessing the results please recall that the Mach number regimes and model geometries differ considerably. Selection of the radius of curvature at the 10% chord location is consistent but arbitrary, although it does seem representative for most blades and gives a good fit for the results. Measured spanwise wavelengths of the periodic vortex arrays on blading are predicted well by the Kestin and Wood theory. If this behavior is at all common it could have implications for turbine aerodynamic and blade cooling design. The outcome is to establish that organized streamwise vorticity may occur more frequently on convex surfaces, such as turbine blade suction surfaces, than hitherto appreciated. Investigations and predictions of flow behavior should be extended to encompass that possibility.

Description: Active Flow Control increases the permissible aerodynamic loading. Curved surface near the trailing edge ("Coanda surface"): a) increases turning -〉 higher pressure ratio. b) controls boundary layer separation -〉 increased surge margin. Objective: Reduce the number of vanes or compressor stages. Constraints: 1. In a real compressor, the vane must still function entirely without blowing. 2. Maintain the flow exit angle of the reference stator despite the resulting increase in stator loading.

Description: Effects of roughness in boundary layers have to be addressed. Until adverse pressure gradient effects are understood, roughness will not significantly drive design. Mechanisms responsible for separation not understood. Effects on Zero Pressure Gradient boundary layers (shear stress). Effects on separation in pressure gradient (prediction of separation). Effect on scalar transport (heat transfer) not understood. Model for skin friction needed in simulations - first grid point likely to be in buffer layer. Definition of roughness important for useful experiments. A lot of validation experiments will be needed. How to get to ks for roughness of engineering interest? - depends on wavelength height, etc. for engineering interest? Re-discovering the wheel should be avoided: existing knowledge (theoretical and experimental) should find its way into the engineering models. It is a task of the industry to filter out the existing information in the literature for results relevant to its application, being external or internal.

Description: Modern turbofan engines employ a highly loaded fan stage with transonic or low-supersonic velocities in the blade-tip region. The fan blades are often prone to flutter at off-design conditions. Flutter is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high-cycle fatigue blade failure. The origins of blade flutter are not fully understood yet. Experimental data that can be used to clarify the origins of blade flutter in modern transonic fan designs are very limited. The Transonic Flutter Cascade Facility at the NASA Glenn Research Center was developed to experimentally study the details of flow mechanisms associated with fan flutter. The cascade airfoils are instrumented to measure high-frequency unsteady flow variations in addition to the steady flow data normally recorded in cascade tests. The test program measures the variation in surface pressure in response to the oscillation of one or more of the cascade airfoils. However, during the initial phases of the program when all airfoils were in fixed positions, conditions were found where significant time variations in the pressures near the airfoil leading edges could be observed.

Description: The NASA Glenn Research Center and Lockheed Martin Corporation tested an aircraft model in two wind tunnels to compare low-speed (subsonic) flow characteristics. Objectives of the test were to determine and document the similarities and uniqueness of the tunnels and to validate that Glenn's 10- by 10-Foot Supersonic Wind Tunnel (10x10 SWT) is a viable low-speed test facility. Results from two of Glenn's wind tunnels compare very favorably and show that the 10x10 SWT is a viable low-speed wind tunnel. The Subsonic Comparison Test was a joint effort by NASA and Lockheed Martin using the Lockheed Martin's Joint Strike Fighter Concept Demonstration Aircraft model. Although Glenn's 10310 and 836 SWT's have many similarities, they also have unique characteristics. Therefore, test data were collected for multiple model configurations at various vertical locations in the test section, starting at the test section centerline and extending into the ceiling and floor boundary layers.

Description: A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

Description: High performance aircraft are, by their very nature, often required to undergo maneuvers involving high angles of attack. Under these conditions unsteady vortices emanating from the wing and the fuselage will impinge on the twin fins (required for directional stability) causing excessive buffet loads, in some circumstances, to be applied to the aircraft. These loads result in oscillatory stresses, which may cause significant amounts of fatigue damage. Active control is a possible solution to this important problem. A full-scale test was carried out on an F/A-18 fuselage and fins using piezoceramic actuators to control the vibrations. Buffet loads were simulated using very powerful electromagnetic shakers. The first phase of this test was concerned with the open loop system identification whereas the second stage involved implementing linear time invariant control laws. This paper looks at some of the problems encountered as well as the corresponding solutions and some results. It is expected that flight trials of a similar control system to alleviate buffet will occur as early as 2001.

Description: Efforts to reduce viscous drag on airfoils could results in a considerable saving for the operation of flight vehicles including those of space transportation. This reduction of viscous drag effort requires measurement and active control of boundary layer flow property on an airfoil. Measurement of viscous drag of the boundary layer flow over an airfoil with minimal flow disturbance is achievable with newly developed MEMS sensor clusters. These sensor clusters provide information that can be used to actively control actuators to obtain desired flow properties or design a vehicle to satisfy particular boundary layer flow criteria. A series of MEMS sensor clusters has been developed with a data acquisition and control module for local measurements of shear stress, pressure, and temperature on an airfoil. The sensor cluster consists of two shear stress sensors, two pressure sensors, and two temperature sensors on a surface area of 1.24 mm x 1.86 mm. Each sensor is 300 microns square and is placed on a flexible polyimide sheet. The shear stress sensor is a polysilicon hot-film resistor, which is insulated by a vacuum cavity of 200 x 200 x 2 microns. The pressure sensors are silicon piezoresistive type, and the temperature sensors are also hot film polysilicon resistors. The total size of the cluster including sensors and electrical leads is 1 Omm x 1 Omm x 0.1 mm. A typical sensitivity of shear stress sensor is 150 mV/Pascal, the pressure sensors are an absolute type with a measurement range from 9 to 36 psia with 0.8mV/V/psi sensitivity, and the temperature sensors have a measurement resolution of 0.1 degree C. The sensor clusters are interfaced to a data acquisition and control module that consists of two custom ASICs (Application Specific Integrated Circuits) and a micro-controller. The data acquisition and control module transfers data to a host PC that configures and controls a total of three sensor clusters. Functionality of the entire system has been tested in the laboratory, and preliminary test results are presented.

Description: In commercial aviation, over 30-percent of all fatal accidents worldwide are categorized as Controlled Flight Into Terrain (CFIT) accidents where a fully functioning airplane is inadvertently flown into the ground, water, or an obstacle. An experiment was conducted at NASA Langley Research Center investigating the presentation of a synthetic terrain database scene to the pilot on a Primary Flight Display (PFD). The major hypothesis for the experiment is that a synthetic vision system (SVS) will improve the pilot s ability to detect and avoid a potential CFIT compared to conventional flight instrumentation. All display conditions, including the baseline, contained a Terrain Awareness and Warning System (TAWS) and Vertical Situation Display (VSD) enhanced Navigation Display (ND). Sixteen pilots each flew 22 approach - departure maneuvers in Instrument Meteorological Conditions (IMC) to the terrain challenged Eagle County Regional Airport (EGE) in Colorado. For the final run, the flight guidance cues were altered such that the departure path went into the terrain. All pilots with a SVS enhanced PFD (12 of 16 pilots) noticed and avoided the potential CFIT situation. All of the pilots who flew the anomaly with the baseline display configuration (which included a TAWS and VSD enhanced ND) had a CFIT event.

Description: In this poster, we describe a web-based tool for verification and automatic generation of user interfaces. The verification component of the tool accepts as input a model of a machine and a model of its interface, and checks that the interface is adequate (correct). The generation component of the tool accepts a model of a given machine and the user's task, and then generates a correct and succinct interface. This write-up will demonstrate the usefulness of the tool by verifying the correctness of a user interface to a flight-control system. The poster will include two more examples of using the tool: verification of the interface to an espresso machine, and automatic generation of a succinct interface to a large hypothetical machine.

Description: To achieve NASA's ambitious mission objectives for the future, aircraft and spacecraft will need intelligence to take the correct action in a variety of circumstances. Vehicle intelligence can be defined as the ability to "do the right thing" when faced with a complex decision-making situation. It will be necessary to implement integrated autonomous operations and low-level adaptive flight control technologies to direct actions that enhance the safety and success of complex missions despite component failures, degraded performance, operator errors, and environment uncertainty. This paper will describe the array of technologies required to meet these complex objectives. This includes the integration of high-level reasoning and autonomous capabilities with multiple subsystem controllers for robust performance. Future intelligent systems will use models of the system, its environment, and other intelligent agents with which it interacts. They will also require planners, reasoning engines, and adaptive controllers that can recommend or execute commands enabling the system to respond intelligently. The presentation will also address the development of highly dependable software, which is a key component to ensure the reliability of intelligent systems.

Description: Six direct numerical simulations of turbulent time-evolving strained plane wakes have been examined to investigate the response of a wake to successive irrotational plane strains of opposite sign. The orientation of the applied strain field has been selected so that the flow is the time-developing analogue of a spatially developing wake evolving in the presence of either a favourable or an adverse streamwise pressure gradient. The magnitude of the applied strain rate a is constant in time t until the total strain e(sup at) reaches about four. At this point, a new simulation is begun with the sign of the applied strain being reversed (the original simulation is continued as well). When the total strain is reduced back to its original value of one, yet another simulation is begun with the sign of the strain being reversed again back to its original sign. This process is done for both initially "favourable" and initially "adverse" strains, providing simulations for each of these strain types from three different initial conditions. The evolution of the wake mean velocity deficit and width is found to be very similar for all the adversely strained cases, with both measures rapidly achieving exponential growth at the rate associated with the cross-stream expansive strain e(sup at). In the "favourably" strained cases, the wake widths approach a constant and the velocity deficits ultimately decay rapidly as e(sup -2at). Although all three of these cases do exhibit the same asymptotic exponential behaviour, the time required to achieve this is longer for the cases that have been previously adversely strained (by at approx. equals 1). These simulations confirm the generality of the conclusions drawn in Rogers (2002) regarding the response of plane wakes to strain. The evolution of strained wakes is not consistent with the predictions of classical self-similar analysis; a more general equilibrium similarity solution is required to describe the results. At least for the cases considered here, the wake Reynolds number and the ratio of the turbulent kinetic energy to the square of the wake mean velocity deficit are determined nearly entirely by the total strain. For these measures the order in which the strains are applied does not matter and the changes brought about by the strain are nearly reversible. The wake mean velocity deficit and width, on the other hand, differ by about a factor of three when the total strain returns to one, depending on whether the wake was first "favourably" or "adversely" strained. The strain history is important for predicting the evolution of these quantities.

Description: The greatest efficiency for a lifting surface at supersonic speeds, according to the theoretical considerations of reference 1, can be attained if the leading edge is swept well behind the Mach cone and the highest aspect ratio which is structurally possible is employed. Such a wing, designed for a Mach number of 3.0, would have 80 deg. of sweepback. Aeroelastic effects have 〈 been shown 3 to be considerable for a wing with 60deg of sweepback and designed for a Mach number of 2.0. The wing shown was found theoretically to have considerable loss in maximum lift-drag ratio attributable to aeroelasticity. This wing has 12-per cent-thick Clark-Y airfoils normal to the wing leading edge. If it were of solid aluminum and flying at a dynamic pressure of 2,400 lbs./sq.ft. (flexibility parameter qb(exp. 4) /El(0) = 7.8), analysis indicates that the wing would deflect so as to reduce the maximum lift-drag ratio about 30 per cent.

Description: Three-dimensional transonic flow over a delta wing is investigated using several turbulence models. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition aft of the leading edge or are fully turbulent are performed. These computations show that grid resolution, transition location and turbulence model significantly affect the 3D flowfield.

Description: NASA is currently investigating a new concept of operations for the National Airspace System, designed to improve capacity while maintaining or improving current levels of safety. This concept, known as Distributed Air/Ground Traffic Management (DAG-TM), allows appropriately equipped autonomous aircraft to maneuver freely for flight optimization while resolving conflicts with other traffic and staying out of special use airspace and hazardous weather. While Airborne Separation Assurance System (ASAS) tools would normally allow pilots to resolve conflicts before they become hazardous, evaluation of system performance in sudden, near-term conflicts is needed in order to determine concept feasibility. If an acceptable safety level can be demonstrated in these situations, then operations may be conducted with lower separation minimums. An experiment was conducted in NASA Langley s Air Traffic Operations Lab to address issues associated with resolving near-term conflicts and the potential use of lower separation minimums. Sixteen commercial airline pilots flew a total of 32 traffic scenarios that required them to use prototype ASAS tools to resolve close range pop-up conflicts. Required separation standards were set at either 3 or 5 NM lateral spacing, with 1000 ft vertical separation being used for both cases. Reducing the lateral separation from 5 to 3 NM did not appear to increase operational risk, as indicated by the proximity to the intruder aircraft. Pilots performed better when they followed tactical guidance cues provided by ASAS than when they didn't follow the guidance. As air-air separation concepts are evolved, further studies will consider integration issues between ASAS and existing Airborne Collision Avoidance Systems (ACAS).These types of non-normal events will require the ASAS to provide effective alerts and resolutions prior to the time that an Airborne Collision Avoidance System (ACAS) would give a Resolution Advisory (RA). When an RA is issued, a pilot must take immediate action in order to avoid a potential near miss. The Traffic Alert and Collision Avoidance System (TCAS) II currently functions as an ACAS aboard commercial aircraft. Depending on the own aircraft s altitude, TCAS only issues RA s 15-35 seconds prior to the Closest Point of Approach (CPA). Prior to an RA, DAG-TM pilots operating autonomous aircraft must rely solely on ASAS for resolution guidance. An additional area of DAG-TM concept feasibility relates to a potential reduction in separation standards. Lower separation standards are likely needed in order to improve NAS efficiency and capacity. Current separation minimums are based in large part on the capabilities of older radar systems. Safety assessments are needed to determine the feasibility of reduced separation minimums. They will give strong consideration to surveillance system performance, including accuracy, integrity, and availability. Candidate surveillance systems include Automatic Dependent Surveillance-Broadcast (ADS-B) and multi-lateration systems. Considering studies done for Reduced Vertical Separation Minimums (RVSM) operations, it is likely that flight technical errors will also be considered. In addition to a thorough evaluation of surveillance system performance, a potential decision to lower the separation standards should also take operational considerations into account. An ASAS Safety Assessment study identified improper maneuvering in response to a conflict (due to ambiguous or improper resolution commands or a pilot s failure to comply with the resolution) as a potential safety risk. If near-term conflicts with lower separation minimums were determined to be more challenging for pilots, the severity of these risks could be even greater.

Description: The placement of actuators on a wing determines the control effectiveness of the airplane. One approach to placement maximizes the moments about the pitch, roll, and yaw axes, while minimizing the coupling. For example, the desired actuators produce a pure roll moment without at the same time causing much pitch or yaw. For a typical wing, there is a large set of candidate locations for placing actuators, resulting in a substantially larger number of combinations to examine in order to find an optimum placement satisfying the mission requirements and mission constraints. A genetic algorithm has been developed for finding the best placement for four actuators to produce an uncoupled pitch moment. The genetic algorithm has been extended to find the minimum number of actuators required to provide uncoupled pitch, roll, and yaw control. A simplified, untapered, unswept wing is the model for each application.

Description: Flow separation in the low-pressure turbine (LPT) is a major barrier that limits further improvements of aerodynamic designs of turbine airfoils. The separation is responsible for performance degradation, and it prevents the design of highly loaded airfoils. The separation can be delayed, reduced, or eliminated completely if flow control techniques are used. Successful flow control technology will enable breakthrough improvements in gas turbine performance and design. The focus of this research project was the development and experimental demonstration of active separation control using glow discharge plasma (GDP) actuators in flow conditions simulating the LPT. The separation delay was shown to be successful, laying the foundation for further development of the technologies to practical application in the LPT. In a fluid mechanics context, the term "flow control" means a technology by which a very small input results in a very large effect on the flow. In this project, the interest is to eliminate or delay flow separation on LPT airfoils by using an active flow control approach, in which disturbances are dynamically inserted into the flow, they interact with the flow, and they delay separation. The disturbances can be inserted using a localized, externally powered, actuating device, examples are acoustic, pneumatic, or mechanical devices that generate vibrations, flow oscillations, or pulses. A variety of flow control devices have been demonstrated in recent years in the context of the external aerodynamics of aircraft wings and airframes, where the incoming flow is quiescent or of a very low turbulence level. However, the flow conditions in the LPT are significantly different because there are high levels of disturbances in the incoming flow that are characterized by high free-stream turbulence intensity. In addition, the Reynolds number, which characterizes the viscous forces in the flow and is related to the flow speed, is very low in the LPT passages.

Description: Researchers are constantly pursuing technologies that will increase the performance of gas turbine engines. The aspirated compressor concept discussed here would allow the compression system to perform its task with about one-half of the compressor blades. To accomplish this, the researchers applied boundary layer control to the blades, casing, and hub. This method of boundary layer control consisted of removing small amounts of air from the main flow path at critical areas of the compressor. This bleed air could be used by other systems such as engine cooling or could be re-injected into lower pressure areas that require air for enhanced performance. This effort was initiated by the Massachusetts Institute of Technology (MIT) in response to a solicitation from the Defense Advanced Research Projects Agency (DARPA) who sought to advance research in flow control technology. The NASA Glenn Research Center partnered with MIT (principal investigator), Honeywell Aircraft Engines (cycle analysis, structural analysis, and mechanical design), and Pratt & Whitney (cycle analysis and aero-analysis) to conceptualize, design, analyze, build, and test the aspirated fan stage. The aero-design and aero-analysis of this fan stage were jointly executed by MIT and Glenn to minimize the amount of bleed flow needed and to maintain the highest efficiency possible (ref. 1). Mechanical design issues were complicated by the need to have a shrouded rotor with hollow blades, with rotor stress levels beyond the capabilities of titanium. The high stress issues were addressed by designing a shroud that was filament wound with a carbon fiber/epoxy matrix, resulting in an assembly that was strong enough to handle the high stresses. Both the rotor (preceding photographs) and stator (following photograph) were fabricated in two halves and then bolted together at the hub and tip, permitting the bleed passages to be machined into each half before assembly.

Description: During the life of any gas turbine, blade erosion is present, especially for those units that are exposed to unfiltered air, such as aviation turbofan engines. The effect of this erosion is to reduce the blade chord progressively from the midspan to the tip region and to roughen and distort the blade surface. The effects of roughness on rotor performance have been documented by Suder et al. and Roberts. These papers indicate that the penalty for leading-edge roughness and erosion can be significant. Turbofan operators, therefore, restore chord length at routine maintenance intervals to regain performance before deterioration is too severe to salvage blades. As the rotor blades erode, the leading edge becomes rough - blunt and distorted from the nominal shape - and the aerodynamic performance suffers. Nominal performance can be recovered by recontouring the leading edges. This process, which inherently shortens the blade chord, can be used until the blade chord erodes to the stall limit. Below this chord length, which varies among engine-compressor types, a decrease of stall margin is likely. After compressor blade rework that includes leading edge recontouring, the blades have different chord lengths, ranging from blades that are near nominal chord length down to those near the stall chord limit. Furthermore, as blades erode below the stall limit, they must be replaced with new blades that have the full nominal chord length. Consequently, a set of compressor blades with varying chord lengths will be installed into each turbofan engine that goes through a complete maintenance cycle. The question arises, "Does fan or compressor performance depend on the order in which mixed-chord blades are installed into a fan or compressor disk?"

Description: Challenges to computational aerothermodynamic (CA) simulation and validation of hypersonic flow over planetary entry vehicles are discussed. Entry, descent, and landing (EDL) of high mass to Mars is a significant driver of new simulation requirements. These requirements include simulation of large deployable, flexible structures and interactions with reaction control system (RCS) and retro-thruster jets. Simulation of radiation and ablation coupled to the flow solver continues to be a high priority for planetary entry analyses, especially for return to Earth and outer planet missions. Three research areas addressing these challenges are emphasized. The first addresses the need to obtain accurate heating on unstructured tetrahedral grid systems to take advantage of flexibility in grid generation and grid adaptation. A multi-dimensional inviscid flux reconstruction algorithm is defined that is oriented with local flow topology as opposed to grid. The second addresses coupling of radiation and ablation to the hypersonic flow solver--flight- and ground-based data are used to provide limited validation of these multi-physics simulations. The third addresses the challenges of retro-propulsion simulation and the criticality of grid adaptation in this application. The evolution of CA to become a tool for innovation of EDL systems requires a successful resolution of these challenges.

Description: The Electric Arc Shock Tube (EAST) facility and the Hypervelocity Free Flight Aerodynamic Facility (HFFAF) at NASA Ames Research Center are described. These facilities have been in operation since the 1960s and have supported many NASA missions and technology development initiatives. The facilities have world-unique capabilities that enable experimental studies of real-gas aerothermal, gas dynamic, and kinetic phenomena of atmospheric entry.

Description: This presentation reviews the three pillars and the associated goals of NASA's Aero-Space Technology Enterprise. The three pillars for success are: (1) Global Civil Aviation, (2) Revolutionary Technology Leaps, (3) Advanced Space Transportation. The associated goals of the first pillar are to reduce accidents, emissions, and cost, and to increase the aviation system capacity. The goals of the second pillar are to reduce transoceanic travel time, revolutionize general aviation aircraft, and improve development capacity. The goals associated with the third pillar are to reduce the launch cost for low earth orbit and to reduce travel time for planetary missions. In order to meet these goals NASA must provide next-generation design capability for new and or experimental craft which enable a balance between reducing components of the design cycle by up to 50% and or increasing the confidence in design by 50%. These next-generation design tools, concepts, and processes will revolutionize vehicle development. The presentation finally reviews the importance of modeling and simulation in achieving the goals.

Description: The supercritical airfoil was chosen as a relatively modem airfoil for comparison. The BOO12 model was tested first. Three different types of flutter instability boundaries were encountered, a classical flutter boundary, a transonic stall flutter boundary at angle of attack, and a plunge instability near M = 0.9 and for zero angle of attack. This test was made in air and was Transonic Dynamics Tunnel (TDT) Test 468. The BSCW model (for Benchmark SuperCritical Wing) was tested next as TDT Test 470. It was tested using both with air and a heavy gas, R-12, as a test medium. The effect of a transition strip on flutter was evaluated in air. The B64AOlO model was subsequently tested as TDT Test 493. Some further analysis of the experimental data for the BOO12 wing is presented. Transonic calculations using the parameters for the BOO12 wing in a two-dimensional typical section flutter analysis are given. These data are supplemented with data from the Benchmark Active Controls Technology model (BACT) given and in the next chapter of this document. The BACT model was of the same planform and airfoil as the BOO12 model, but with spoilers and a trailing edge control. It was tested in the heavy gas R-12, and was instrumented mostly at the 60 per cent span. The flutter data obtained on PAPA and the static aerodynamic test cases from BACT serve as additional data for the BOO12 model. All three types of flutter are included in the BACT Test Cases. In this report several test cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flutter. Cases are selected for classical and stall flutter for the BSCW model, for classical and plunge for the B64AOlO model, and for classical flutter for the BOO12 model. Test Cases are also presented for BSCW for static angles of attack. Only the mean pressures and the real and imaginary parts of the first harmonic of the pressures are included in the data for the test cases, but digitized time histories have been archived. The data for the test cases are available as separate electronic files. An overview of the model and tests is given, the standard formulary for these data is listed, and some sample results are presented.

Description: The Efficient Descent Advisor (EDA) controller automation tool generates trajectory-based speed, path, and altitude-profile advisories to facilitate efficient, continuous descents into congested terminal airspace. While prior field trials have assessed the trajectory-prediction accuracy for large jet (i.e., Boeing and Airbus) types, smaller (i.e., regional and business) jet types present unique challenges involving different descent procedures and Flight Management System (FMS) capabilities. A small-jet field trial was conducted at Denver in the fall of 2010 with the objective of measuring trajectory prediction accuracy and quantifying the primary sources of error. This paper uses data collected onboard a Bombardier Global 5000 test aircraft to quantify the size and sources of trajectory prediction error. Error sources were quantified for the 44 runs by incrementally replacing predicted data with data collected onboard the aircraft and measuring the effect on time error. Results for en-route descents, from prior to top of descent to the meter fix 60-120 nmi downstream, indicate that the aircraft arrived an average 15 seconds earlier than predicted, with a standard deviation of 10 seconds. Target Mach and CAS deceleration were found to be the two largest error sources. If CAS deceleration error was reduced using a typical, more predictable level flight deceleration then the arrival time prediction error in 2010 would be on par with a 2009 flight trial of Airbus and Boeing revenue flights. Four of the error sources, tracker jumps, CAS deceleration, target Mach, and path distance, lend themselves to significant reductions with modest to no changes to ATC automation andor procedures. Wind error and its impact on arrival time error was significantly reduced in 2010 compared to a 1994 flight test using NASAs Boeing 737 test aircraft.

Description: A simplified method is presented for estimating the lift-curve slope of irregular planform wings at subsonic speeds and low angles of attack. The present process is an extension of the method derived in NACA Technical Note 3911 and enables quick estimates of subsonic liftcurve slope, to be made whereas more refined procedures require considerable time and computation. Comparison of experimental and estimated values for a wide range of wing planforms having discontinuous spanwise sweep variation indicates good agreement. A comparison of the present procedure with a 20-step vortex method (NACA Research Memorandum L50L13) indicated good agreement for a variable-sweep configuration.

Description: No abstract available

Description: An analysis is presented of the influence of wing aspect ratio and tail location on the effects of compressibility upon static longitudinal stability. The investigation showed that the use of reduced wing aspect ratios or short tail lengths leads to serious reductions in high-speed stability and the possibility of high-speed instability.

Description: Pressure distribution over an extended leading-edge flap on a 42 degree swept-back wing was investigated. Results indicate that the flap normal-force coefficient increased almost linearly with the angle of attack to a maximum value of 3.25. The maximum section normal-force coefficient was located about 30 percent of the flap span outboard of the inboard end and had a value of 3.75. Peak negative pressures built up at the flap leading edge as the angle of attack was increased and caused the chordwise location of the flap center of pressure to be move forward.

Description: Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.

Description: An experimental investigation was made of a preloaded spring-tab flutter model to determine the effects on flutter speed of aspect ratio, tab frequency, and preloaded spring constant. The rudder was mass-balanced, and the flutter mode studied was essentially one of three degrees of freedom (fin bending coupled with rudder and tab oscillations). Inasmuch as the spring was preloaded, the tab-spring system was a nonlinear one. Frequency of the tab was the most significant parameter in this study, and an increase in flutter speed with increasing frequency is indicated. At a given frequency, the tab of high aspect ratio is shown to have a slightly lower flutter speed than the one of low aspect ratio. Because the frequency of the preloaded spring tab was found to vary radically with amplitude, the flutter speed decreased with increase in initial displacement of the tab.

Description: In the course of a flight test of a supersonic research pilotless aircraft (the NACA RM-1), large-amplitude aileron oscillations, probably aileron compressibility flutter, were encountered in the transonic and supersonic speed ranges. The wing was oscillating at the same frequency as the aileron. The aircraft was equipped with 45 degree swept-back wings of symmetrical NASA 65-010 airfoil section. Completely mass-balanced ailerons with 20 degree beveled trailing edges were installed on the wings. The ailerons were free floating with no mechanical restraining force other than the friction of the aileron hinges and servomechanism bearings throughout the high-speed interval of flight.

Description: A three-dimensional investigation of straight-sided-profile plain ailerons on a wing with 30 degrees and 45 degrees of sweepback and sweepforward was made in a high-speed wind tunnel for aileron deflections from -10 degrees to 10 degrees and at Mach numbers from 0.60 to 0.96. Wing configurations of 30 degrees generally reduced the severity of the large changes in rolling-moment and aileron hinge-moment coefficients experienced by the upswept wing configurations as the result of compression shock and extended to higher Mach numbers the speeds at which such changes occurred.

Description: Low-speed tests of a pilotless aircraft were conducted in the Langley propeller-research tunnel to provide information for the estimation of the longitudinal stability and. control, to measure the aileron effectiveness, and to calibrate the radome and the Machmeter pitot-static orifices. It was found that the model possessed a stEb.le variation of elevator angle required for trim throughout the speed range at the design angle of attack. A comparison of the airplane with and without JATO units and with an alternate rocket booster showed that a large loss in longitudinal stability and control resulting from the addition of the rocket booster to the aircraft was sufficient to make the rocket-booster assembly unsatisfactory as an alternate for the JATO units. Reversal of the aileron effectiveness was evident at positive deflections of the vertical wing flap indicating that the roll-stabilization system would produce roiling moments in a tight right turn contrary to its design purpose. Vertical-wing-flap deflections caused large errors in the static-pressure reading obtained by the original static-tube installation. A practical installation point on the fuselage was located which should yield reliable measurement of the free-stream static pressure.

Description: On the basis of a recently developed theory for finite sweptback wings at supersonic speeds, calculations of the supersonic wave drag at zero lift were made for a series of wings having thin symmetrical biconvex sections with untapered plan forms and various angles of sweepback and aspect ratios. The results are presented in a unified form so that a single chart permits the direct determination of the wave drag for this family of airfoils for an extensive range of aspect ratio and sweepback angle for stream Mach numbers up to a value corresponding to that at which the Mach line coincides with the wing leading edge. The calculations showed that in general the wave-drag coefficient decreased with increasing sweepback. At Mach numbers for which the Mach lines are appreciably ahead of the wing leading edge, the 'wave-drag coefficient decreased to an important extent with increases in aspect ratio or slenderness ratio. At Mach numbers for which the Mach lines approach the wing leading edge (Mach numbers approaching a value equal to the secant of the angle of sweepback), the wave-drag coefficient decreased with reductions in aspect ratio or slenderness ratio. In order to check the results obtained by the theory, a comparison was made with the results of tests at the Langley Memorial Aeronautical Laboratory of sweptback wing attached to a freely falling body. The variation of the drag with Mach number and aspect ratio as given by the theory appeared to be in reasonable

Description: No abstract available

Description: An investigation has been conducted in the Cleveland 18- by 18-inch supersonic tunnel at a Mach number of 1.85 and angles of attack from 0 deg to 5 deg to determine optimum design configurations for a convergent-divergent type of supersonic diffuser with a subsonic diffuser of 5 deg included divergence angle. Total pressure recoveries in excess of theoretical recovery across a normal shock at a free-stream Mach number of 1.85 wore obtained with several configurations. The highest recovery for configurations without a cylindrical throat section was obtained with an inlet having an included convergence angle of 20 deg. Insertion of a 2-inch throat section between a 10 deg included angle inlet and the subsonic diffuser stabilized the shock inside the diffuser and resulted in recoveries as high as 0.838 free-stream total pressure at an angle of attack of 0 deg, corresponding to recovery of 92.4 percent of the kinetic energy of the free air stream. Use of the throat section also lessened the reduction in recovery of all configurations due to angle of attack.

Description: A wind-tunnel investigation has been made to determine the effects of unsymmetrical horizontal-tail arrangements on the power-on static longitudinal stability of a single-engine single-rotation airplane model. Although the tests and analyses showed that extreme asymmetry in the horizontal tail indicated a reduction in power effects on longitudinal stability for single-engine single-rotation airplanes, the particular "practical" arrangement tested did not show marked improvement. Differences in average downwash between the normal tail arrangement and various other tail arrangements estimated from computed values of propeller-slipstream rotation agreed with values estimated from pitching-moment test data for the flaps-up condition (low thrust and torque) and disagreed for the flaps-down condition (high thrust and torque). This disagreement indicated the necessity for continued research to determine the characteristics of the slip-stream behind various propeller-fuselage-wing combinations. Out-of-trim lateral forces and moments of the unsymmetrical tail arrangements that were best from consideration of longitudinal stability were no greater than those of the normal tail arrangement.

Description: Wing was tested with full-span, partial-span, or split flaps deflected 60 Degrees and without flaps. Chordwise pressure-distribution measurements were made for all flap configurations.. Peak values of maximum lift coefficient were obtained at relatively low free-stream Mach numbers and, before critical Mach number was reached, were almost entirely dependent on Reynolds Number. Lift coefficient increased by increasing Mach number or deflecting flaps while critical pressure coefficient was reached at lower free-stream Mach numbers.

Description: Theoretical analysts of lateral dynamic motion of tailless and conventional airplanes was made for fighter and heavy transport. Their reactions to a lateral gust and control power required by each for simple maneuvers were determined and compared. Both types of airplanes require almost identical aileron control power to perform a given maneuver; tailless airplane requires about 1-2 to 1-3 directional control power of conventional airplane. Tailless airplane also shows greatest displacement for a given disturbance and has least damping in oscillatory mode.

Description: For the normal range of engine power the impeller provided marked improvement over the standard spray-bar injection system. Mixture distribution at cruising was excellent, maximum cylinder temperatures were reduced about 30 degrees F, and general temperature distribution was improved. The uniform mixture distribution restored the normal response of cylinder temperature to mixture enrichment and it reduced the possibility of carburetor icing, while no serious loss in supercharger pressure rise resulted from injection of fuel near the impeller outlet. The injection impeller also furnished a convenient means of adding water to the charge mixture for internal cooling.

Description: Behaviors of both model and full-scale airplanes were ascertained by making visual observations, by recording time histories of decelerations, and by taking motion picture records of ditchings. Results are presented in form of sequence photographs and time-history curves for attitudes, vertical and horizontal displacements, and longitudinal decelerations. Time-history curves for attitudes and horizontal and vertical displacements for model and full-scale tests were in agreement; maximum longitudinal decelerations for both ditchings did not occur at same part of run; full-scale maximum deceleration was 50 percent greater.

Description: Finite trigonometric series is fitted by harmonic analysis as an approximation function to the psi function of the Theodorsen arbitrary-airfoil potential theory. By harmonic synthesis, the corresponding conjugate trigonometric series is used as an approximation to the epsilon function. A set of coefficients of particularly simple form is obtained algebraically for direct calculation of the epsilon values from the corresponding set of psi values. Complete derivation of this process is presented.

Description: A program of model tests has been completed at Langley tank no. 1 which will furnish a qualitative guide as to the relation of length of afterbody and depth of step. The model used for the tests was a l/12-size unpowered dynamic model of a hypothetical 160,000-pound airplane. The results showed that an increase in length of afterbody requires an accompanying increase in depth of step to maintain adequate landing stability. Changing the length of afterbody and depth of step in such a manner as to maintain a given landing stability will result in only small changes in take-off stability.

Description: Tests show that at inlet-air temperatures of 250 deg F and 100 deg F the knock-limited performance of the base fuel of blends, leaded with 4 ml TEL per gallon and containing 20 percent spiropentane, was reduced at fuel/air ratios below 0.085. The 20 percent methylenecyclobutane reduced the knock-limited power of the base fuel at fuel/air ratios below 0.112. Di-tert-butyl ether, methyl-tert-butyl ether, and triptane increased the knock-limited power of the base fuel at all fuel/air ratios and at both temperatures.

Description: Engine temperature data and cooling correlating analyses of the engine and oil cooler are presented in connection with an investigation of the cowling and cooling of the ranger V-770-8 engine installation in the Edo XOSE-1 airplane. Three types of baffles were installed in the course of the tests: the conventional, the turbulent-flow, and the NACA diffuser baffles. Each of the types was of merit in cooling a different region on the cylinder. Incorporation of the best features of the three types into one baffle, a method which appears to be feasible, would provide improvements in cylinder cooling.

Description: Lift characteristics and pressure distribution for a NACA 230 wing were investigated for an angle of attack range of from -10 to +24 degrees and Mach range of from 0.2 to 0.7. Maximum lift coefficient increased up to a Mach number of 0.3, decreased rapidly to a Mach number of 0.55, and then decreased moderately. At high speeds, maximum lift coefficient was reached at from 10 to 12 degrees beyond the stalling angle. In high-speed stalls, resultant load underwent a moderate shift outward.

Description: Sectional characteristics of airfoil having retractable slotted flap with plain, slot-lip, or retractable ailerons are presented for a large range of aileron deflections. The analysis indicated that pitching moments produced by spoilers were less positive than those produced by plain flaps of equal effectiveness, also that pitching moments created by the spoiler increased less with the Mach number than similar moments produced by plain flaps. Positive values of pitching moment decreased as devices were located nearer airfoil leading edge.

Description: Results of flight tests of a control-feel aid presented. This device consisted of a spring and dashpot connected in series between the control stick and airplane structure. The device was tested in combination with an experimental elevator and bobweight which had given unsatisfactory dynamic stability and control-feel characteristics in previous tests. The control-feel aid effected marked improvement in both the control-feel characteristics and the control-feel dynamic longitudinal stability of the airplane.

Description: Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined, from two-dimensional windtunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NPiCA 6-series airfoils. The experimental results confirm the design expectations in demonstrating for the NACA S-series airfoils either no variation, or an Increase from the low-speed design value, In the lift coefficient at a constant angle of attack with increasing Mach number above the critical. It was not found possible to improve the variation with Mach number of the slope of the lift curve for these airfoils above that for the NACA 6-series airfoils. The drag characteristics of the new airfoils are somewhat inferior to those of the NACA 6- series with respect to divergence with Mach number, but the pitching-moment characteristics are more favorable for the thinner new sections In demonstrating somewhat smaller variations of moment coefficient with both angle of attack and Mach number. The effect on the aero&ynamic characteristics at high Mach numbers of removing the cusp from the trailing-edge regions of two 10-percent-chord-thick NACA 6-series airfoils is determined to be negligible.

Description: Investigations were made to develop a simplified method for designing exhaust-pipe shrouds to provide desired or maximum cooling of exhaust installations. Analysis of heat exchange and pressure drop of an adequate exhaust-pipe shroud system requires equations for predicting design temperatures and pressure drop on cooling air side of system. Present experiments derive such equations for usual straight annular exhaust-pipe shroud systems for both parallel flow and counter flow. Equations and methods presented are believed to be applicable under certain conditions to the design of shrouds for tail pipes of jet engines.

Description: The first part of this paper reviews the present state of the problem of the instability of laminar boundary layers which has formed an important part of the general lectures by von Karman at the first and fourth Congresses and by Taylor at the fifth Congress. This problem may now be considered as essentially solved as the result of work completed since 1938. When the velocity fluctuations of the free-stream flow are less than 0.1 percent of the mean speed, instability occurs as described by the well-known Tollmien-Schlichting theory. The Tollmien-Schlichting waves were first observed experimentally by Schubauer and Skramstad in 1940. They devised methods of introducing controlled small disturbances and obtained measured values of frequency, damping, and wave length at various Reynolds numbers which agreed well with the theoretical results. Their experimental results were confirmed by Liepmann. Much theoretical work was done in Germany in extending the Tol1mien-Schlichting theory to other boundary conditions, in particular to flow along a porous wall to which suction is applied for removing part of the boundary layer. The second part of this paper summarizes the present state of knowledge of the mechanics of turbulent boundary layers, and of the methods now being used for fundamental studies of the turbulent fluctuations in turbulent boundary layers. A brief review is given of the semi-empirical method of approach as developed by Buri, Gruschwitz, Fediaevsky, and Kalikhman. In recent years the National Advisory.Commsittee for Aeronautics has sponsored a detailed study at the National Bureau of Standards of the turbulent fluctuations in a turbulent boundary layer under adverse pressure gradient sufficient to produce separation. The aims of this investigation and its present status are described.

Description: An experimental investigation was conducted to determine the penetration of a circular air Jet directed perpendicularly to an air stream as a function of Jet density, Jet velocity, air-stream density, air-stream velocity, Jet diameter, and distance downstream from the Jet. The penetration was determined for nearly constant values of air-stream density at two tunnel velocities, four Jet diameters, four positions downstream of the Jet, and for a large range of Jet velocities and densities. An equation for the penetration was obtained in terms of the Jet diameter, the distance downstream from the jet, and the ratios of Jet and air-stream velocities and densities.

Description: Tests of a partial-span model of a large bomber-type air1ane were conducted to determine the. aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans. With full-span flaps deflected and with the 2-percent vent open or closed the initial stalling of the wing occurred at the tips, but with the vents closed there probably would be no appreciable loss in lateral control until maximum lift was reached. The l-percent vent increased the rolling effectiveness of the spoiler at small spoi1er deflections, particularly at high angles of attack with flaps deflected. With flaps deflected the 2-percent vent caused a large reduction in both the wing lift and rolling effectiveness of the spoiler at large angles of attack. However, at small angle of attack the 2-percent vent increased the rolling effectiveness of the spoiler at small spoiler deflections. The simultaneous operation of the spoiler and vent (in contrast to a vent fixed in the wing) would result in a large increase in the effectiveness of the spoiler and would avoid any loss in wing lift as in a fixed vent arrangement. The tests of the spoiler modifications revealed that (1) the spoiler perforations reduced the rolling-moment and yawing-moment coefficients but caused the spoiler hinge-moment coefficients to become more positive; (2) the spoiler slot had no notable effect on the rolling-moment and yawing-moment characteristics but produced a positive increase in the spoiler hinge-moment coefficients at large spoiler deflections; (3) the effects produced by the individual modifications were additive when the various modifications were combined. In general, progressively decreasing the spoiler span by removing the segments from the inboard end of the spoiler caused a decrease in rolling effectiveness approximately proportional to the span of the segment.

Description: A theoretical investigation was conducted on jet-induced flow deviation. Analysis is given of flow inclination induced outside cold and hot jets and jet deflection caused by angle of attack. Applications to computation of effects of jet on longitudinal stability and trim are explained. Effect of jet temperature on flow inclination was found small when thrust coefficient is used as criterion for similitude. The average jet-induced downwash over tail plane was obtained geometrically.

Description: Recent airfoil data for both flight and wind-tunnel tests have been collected and correlated insofar as possible. The flight data consist largely of drag measurements made by the wake-survey method. Most of the data on airfoil section characteristics were obtained in the Langley two-dimensional low-turbulence pressure tunnel. Detail data necessary for the application of NACA 6-serles airfoils to wing design are presented in supplementary figures, together with recent data for the NACA 24-, 44-, and 230-series airfoils. The general methods used to derive the basic thickness forms for NACA 6- and 7-series airfoils and their corresponding pressure distributions are presented. Data and methods are given for rapidly obtaining the approximate pressure distributions for NACA four-digit, five-digit, 6-, and 7-series airfoils. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Available data on high-lift devices are presented. Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed. The data indicate that the effects of surface condition on the lift and drag characteristics are at least as large as the effects of the airfoil shape and must be considered in airfoil selection and the prediction of wing characteristics. Airfoils permitting extensive laminar flow, such as the NACA 6-series airfoils, have much lower drag coefficients at high speed and cruising lift coefficients than earlier types-of airfoils if, and only if, the wing surfaces are sufficiently smooth and fair. The NACA 6-series airfoils also have favorable critical-speed characteristics and do not appear to present unusual problems associated with the application of high-lift and lateral-control devices. Much of the data given in the NACA Advance Confidential Report entitled "Preliminary Low-Drag-Airfoil and Flap Data from Tests at Large Reynolds Number and Low Turbulence," by Eastman N. Jacobs, Ira R. Abbott, and Milton Davidson, March 1942 has been corrected and included in the present paper, which supersedes the previously published paper.

Description: Statistical methods were applied to acceleration and airspeed data obtained with the XC-35 airplane during flights in turbulent air within convective clouds in order to determine the characteristics of repeated or closely spaced gusts pertinent to design problems. Results indicated that, in turbulent air within convective cloud, gusts tend to be contiguous and are seldom found isolated in space. Over-all average spacing between repeated gusts was in good agreement with twice the average gust-gradient distance of 10 chords used in present design.

Description: At the request of the Air Technical Service Command, U.S. Army Air Forces, a 0.22-scale model of a twin-fuselae pursuit airplane was built and tested at the Ames Aeronautical Laboratory. The tests of this model were made in order that the aerodynamic characteristics of the airplane, especially at high speed, might be predicted. The results shown in this report consist of force data for the model and critical Mach numbers of parts of the model as determined from pressure-distribution measurements. The results indicate that a diving tendency of the airplane can be expected at Mach numbers above 0.70 at lift co-efficients from 0 to 0.4. There is an indication that the Mach number at which the airpolane would first experience a diving tendency for lift coefficients from 0 to 0.2 can be increased if the critical speed of the radiator enclosures is increased, and the wing-fuselage-juncture fillets are improved.

Description: Two-dimensional data were obtained in Mach range of from 0.40 to 0.94 and Reynolds Number range of (3.4 - 4.2) X 10 Degrees. Results indicate that thickness ratio is dominating shape parameter at high Mach numbers and that aerodynamic advantages are attainable by using thinnest possible sections. Effects of jet boundaries, Reynolds Number, and Data presented are free from jet-boundary and humidity effects.

Description: Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.

Description: Tests in Langley pressure tunnel of model XA-26 bomber were compared with those of A-26B (twin-engine attack bomber) and showed that static longitudinal stability, indicated by elevator-fixed neutral points, and variation of elevator deflection in straight and turning flight were good. Airplane possessed improved stability at low speeds which was attributed to pronounced stalling at root of production wing. At rudder-force reversal at speeds higher than those in flight tests, agreement in rudder-fixed and rudder-free static directional stability was good. Hinge moment obtained at zero sideslip was satisfactory for determining aileron forces in sideslip.

Description: Closed-form expressions and tables composed from these expressions are presented for complete and partial conic and spheric bodies at combined angles of attack and sideslip in Newtonian flow. Aerodynamic coefficients of these bodies are tabulated for various body segments over a range of angles of attack from 1 deg to 85 deg and angles of sideslip from 0 deg to 15 deg. Some comparisons between Newtonian predictions and hypersonic experimental aerodynamic characteristics were made for conic bodies hawing various surface slopes, nose bluntnesses, and body cross sections to indicate the range of validity of the theory. In general, the theory is shown to agree quite well with experimental results for sharp-nose complete cones and for configurations hawing large blunted noses and steep surface slopes. However, agreement between theory and experiment generally is poor for the more slender, slightly blunted complete or half conic bodies and also for sharp-nose half conic bodies where real-flow phenomena such as forebody interference, viscous forces, leeward surface contributions, or leading-edge pressure reductions may have significant effect. The agreement between theory and experiment for the bodies considered can be improved by using the stagnation pressure coefficient behind a normal shock rather than 2 as the Newtonian coefficient, although for the sharp-nose half conic bodies there i s no theoretical justification for this modification.

Description: An investigation was made of the static longitudinal stability, and control and stall characteristics of XBTK-1 dive bomber. Results indicate that the longitudinal stability will probably be satisfactory for all contemplated flight conditions at the rear-most CG location with elevator both fixed and free. Power effects were small. Sufficient elevator control will be available to trim in any flight condition above the ground. Increasing the slotted flap deflection above 30 degrees only slightly increased the max. lift coefficient.

Description: Availability data obtained on SNB-1 trainer-class airplanes were analyzed and results presented as flight envelopes which predict occurrences of large values of air speed and acceleration. Comparison is made with SNJ-4 trainer-class airplane data analyzed by the same method. It is concluded that flight envelopes are satisfactory; that the two types show large differences in flight loads and speeds experience; and that SNB-1 will seldom, if ever, exceed design limit load factor and restricted speed, which SNJ-4 can be expected to exceed design-limit load factor and restricted speed in a very small number of flight hours.

Description: The data presented have no bearing on performance characteristics of airplane, which were considered exceptionally good in previous tests. Some of the undesirable features of lateral and directional stability and control characteristics of the F-8 are listed. Directional stability, with rudder fixed, did not sufficiently restrict aileron yaw; rudder control was inadequate during take-off and landing, and was insufficient to fly airplane with one engine; in clean condition, power of ailerons was slightly below minimum value specified; it was difficult to trim airplane in rough air.

Description: Results are reported of knock-limited tests of five aromatics, each individually blended with selected base fuels and tested with and without TEL, using 17.6, F-4, and F-3 small-scale engines. The five aromatics rated in the following order of decreasing antiknock effectiveness at fuel/air ratio 0.10: m-xylene, 1-isopropyl-4-methylbenzene, n-propylbenzene, isobutylbenzene, and n-butylbenzene.

Description: Data are presented of the flow conditions in the vicinity of an NACA D sub S -type cowling. Tests were made of a 1/2 scale-nacelle model at inlet-velocity ratios ranging from 0.23 to 1.02 and angles of attack from 6 deg to 10 deg. The velocity and direction of flow in the vertical plane of symmetry of the cowling were determined from orifices and tufts installed on a board aligned with the flow. Diagrams showing velocity ratio contours and lines of constant flow angles are given.

Description: In Mach range of 0.25 - 0.69, boundary-layer measurements were made on upper wing surface at 25% semi-span, pressure-distribution measurements made on upper surface at 63% semi-span, and wake surveys made at 63% semi-span. The minimum profile-drag coefficient of 0.0062 was indicated for smooth section at 63% semi-span. Critical mach number was exceeded by 0.04, but no compressibility shocks appeared. In slipstream, boundary layer transition occurred as far back as 20% chord on upper surface at low lift coefficients.

Description: Lift, drag, internal flow, and pressure distribution measurements were made on a low-drag airfoil incorporating various air inlet designs. Two leading-edge air inlets are developed which feature higher lift coefficients and critical Mach than the basic airfoil. Higher lift coefficients and critical speeds are obtained for leading half of these inlet sections but because of high suction pressures near exist, slightly lower critical speeds are obtained for the entire inlet section than the basic airfoil.

Description: The results of a theoretical analysis of the hinge-moment characteristics of various sealed-internal-balance arrangements for control surfaces are presented. The analysis considered overhands sealed to various types of wing structure by flexible seals spanning gaps of various widths or sealed to the wing structure by a flexible system of linked plates. Leakage was not considered; the seal was assumed to extend the full spanwise length of the control surface. The effect of the developed width of the flexible seal and of the geometry of the structure to which the seal was anchored was investigated, as well as the effect of the gap width that is sealed. The results of the investigation indicated that the most nearly linear control-surface hinge-moment characteristics can probably be obtained from a flexible seal over a narrow gap (about 0.1 of the overhang chord), which is so installed that the motion of the seal is restricted to a region behind the point of attachment of the seal to the wing structure. Control-surface hinge moments that tend to be high at large deflections and low or overbalanced at small deflections will result if a very narrow seal is used.

Description: Tests were made to determine whether spring-tab ailerons tended to oscillate or flutter in speed ranges up to 400 mph. Flight tests showed spring-tab ailerons had desirable light stick forces and no tendency to overbalance. No flutter tendencies were indicated up to 400 mph, and any oscillations following abrupt control deflections were heavily damped. Recommendations were made for modifications to increase aileron effectiveness at low speeds without affecting lateral control at high speeds by increasing available deflection and modifying spring-tab arrangement.

Description: Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.