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Since the project is rapidly nearing conclusion, the status of the tasks outlined in the original proposal are briefly outlined. These tasks include: viscous interation and wake curvature effects; code optimization and design methodology studies; methods for the design of isolated regions; program improvement efforts; and validation, testing, and documentation.
An aerodynamic theory of the flow about slender delta wings is described. The theory includes a treatment of the self-consistent development of the vortex wake patterns above the wing necessary to maintain smooth flow at the wing edges. The paper focuses especially on the formation within the wake of vortex 'cores' as embedded regions of total pressure loss, fed and maintained by umbilical vortex sheets emanating from the wing edges. Criteria are developed for determining the growing size and location of these cores, as well as the distribution and strength of the vorticity within them. In this paper, however, the possibility of vortex breakup is omitted. The aerodynamic consequences of the presence and evolution of the cores and the associated wake structure are illustrated and discussed. It is noted that wake history effects can have substantial influence on the distribution of normal force on the wing as well as on its magnitude.
Computational and experimental performance benefits are presented for a high-aspect-ratio unswept wing configuration with sheared tips. The sheared tip is a highly swept and highly tapered surface located in the same plane as the inboard wing panel to which it is attached. The compuational results were obtained with an inviscid surface panel method that models the nonlinear influence of the trailing wake. Both wind-tunnel and calculated results were obtained for a 12-ft span wing model with various wing-tip configurations. The computational and experimental data are in fair agreement and demonstrate that sheared wing tips can reduce induced drag at cruise and climb lift coefficients. The drag reduction is the result of wake deformation effects and changes in spanwise load distribution. Wind-tunnel measured longitudinal and lateral directional stability characteristics are also presented for the various wing-tip layouts.
Experimental studies of isolated rotors in forward flight have indicated that higher harmonic pitch control can reduce rotor noise. These tests also show that such pitch inputs can generate substantial vibratory loads. The modification is summarized of the RotorCRAFT (Computation of Rotor Aerodynamics in Forward flighT) analysis of isolated rotors to study the vibratory loading generated by high frequency pitch inputs. The original RotorCRAFT code was developed for use in the computation of such loading, and uses a highly refined rotor wake model to facilitate this task. The extended version of RotorCRAFT incorporates a variety of new features including: arbitrary periodic root pitch control; computation of blade stresses and hub loads; improved modeling of near wake unsteady effects; and preliminary implementation of a coupled prediction of rotor airloads and noise. Correlation studies are carried out with existing blade stress and vibratory hub load data to assess the performance of the extended code.
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Many attempts have been made in recent years to predict the off-axis response of a helicopter to control inputs, and most have had little success. Since physical insight is limited by the complexity of numerical simulation models, this paper examines the off-axis response problem using an analytical model, with the goal of understanding the mechanics of the coupling. A new induced velocity model is extended to include the effects of wake distortion from pitch rate. It is shown that the inclusion of these results in a significant change in the lateral flap response to a steady pitch rate. The proposed inflow model is coupled with the full rotor/body dynamics, and comparisons are made between the model and flight test data for a UH-60 in hover. Results show that inclusion of induced velocity variations due to shaft rate improves correlation in the pitch response to lateral cycle inputs.
An experimental/analytical study has been conducted to determine the performance improvements achievable by circumferentially indexing succeeding rows of turbine stator airfoils. A series of tests was conducted to experimentally investigate stator wake clocking effects on the performance of the space shuttle main engine (SSME) alternate turbopump development (ATD) fuel turbine test article (TTA). The results from this study indicate that significant increases in stage efficiency can be attained through application of this airfoil clocking concept. Details of the experiment and its results are documented in part 1 of this paper. In order to gain insight into the mechanisms of the performance improvement, extensive computational fluid dynamics (CFD) simulations were executed. The subject of the present paper is the initial results from the CFD investigation of the configurations and conditions detailed in part 1 of the paper. To characterize the aerodynamic environments in the experimental test series, two-dimensional (2D), time accurate, multistage, viscous analyses were performed at the TTA midspan. Computational analyses for five different circumferential positions of the first stage stator have been completed. Details of the computational procedure and the results are presented. The analytical results verify the experimentally demonstrated performance improvement and are compared with data whenever possible. Predictions of time-averaged turbine efficiencies as well as gas conditions throughout the flow field are presented. An initial understanding of the turbine performance improvement mechanism based on the results from this investigation is described.
Flow computations around two-element and three-element configurations are presented and compared to detailed experimental measurements. The k-epsilon-v(exp 2)(bar) model has been applied and the ability of the model to capture streamline curvature effects, wake-boundary layer confluence, and laminar/turbulent transition is discussed. The numerical results are compared to experimental datasets that include mean quantities (velocity and pressure coefficient) and turbulent quantities (Reynolds normal and shear stresses).
There are more people passing through the world's airports today than at any other time in history. With this increase in civil transport, airports are becoming capacity limited. In order to increase capacity and thus meet the demands of the flying public, the number of runways and number of flights per runway must be increased. In response to the demand, the National Aeronautics and Space Administration (NASA), in conjunction with the Federal Aviation Administration (FAA), airport operators, and the airline industry are taking steps to increase airport capacity without jeopardizing safety. Increasing the production per runway increases the likelihood that an aircraft will encounter the trailing wake-vortex of another aircraft. The hazard of a wake-vortex encounter is that heavy load aircraft can produce high intensity wake turbulence, through the development of its wing-tip vortices. A smaller aircraft following in the wake of the heavy load aircraft will experience redistribution of its aerodynamic load. This creates a safety hazard for the smaller aircraft. Understanding this load redistribution is of great importance, particularly during landing and take-off. In this research wake-vortex effects on an encountering 10% scale model of the B737-100 aircraft are modeled using both strip theory and vortex-lattice modeling methods. The models are then compared to wind tunnel data that was taken in the 30ft x 60ft wind tunnel at NASA Langley Research Center (LaRC). Comparisons are made to determine if the models will have acceptable accuracy when parts of the geometry are removed, such as the horizontal stabilizer and the vertical tail. A sensitivity analysis was also performed to observe how accurately the models could match the experimental data if there was a 10% error in the circulation strength. It was determined that both models show accurate results when the wing, horizontal stabilizer, and vertical tail were a part of the geometry. When the horizontal stabilizer and vertical tail were removed there were difficulties modeling the sideforce coefficient and pitching moment. With the removal of only the vertical tail unacceptable errors occurred when modeling the sideforce coefficient and yawing moment. Lift could not be modeled with either the full geometry or the reduced geometry attempts.
Wake-vortex effects on an 10% scale model of the B737-100 aircraft are calculated using both strip theory and vortex-lattice methods. The results are then compared to data taken in the 30ft x 60ft wind tunnel at NASA Langley Research Center (LaRC). The accuracy of the models for a reduced geometry, such with the horizontal stabilizer and the vertical tail removed, is also investigated. Using a 10% error in the circulation strength and comparing the model's results with the experiment illustrates the sensitivity of the models to the vortex circulation strength. It was determined that both strip theory and the vortex lattice method give accurate results when all the geometrical information is used. When the horizontal stabilizer and vertical tail were removed there were difficulties modeling the sideforce coefficient and pitching moment. With the removal of only the vertical tail unacceptable errors occurred when modeling the sideforce coefficient and yawing moment. Lift could not be accurately modeled with either the full geometry or the reduced geometry.
A series of aerodynamic heating tests was conducted on a 70-deg sphere-cone planetary entry vehicle model in a Mach 10 perfect-gas wind tunnel at freestream Reynolds numbers based on diameter of 8.23x104 to 3.15x105. Surface heating distributions were determined from temperature time-histories measured on the model and on its support sting using thin-film resistance gages. The experimental heating data were compared to computations made using an axisymmetric/2D, laminar, perfect-gas Navier-Stokes solver. Agreement between computational and experimental heating distributions to within, or slightly greater than, the experimental uncertainty was obtained on the forebody and afterbody of the entry vehicle as well as on the sting upstream of the free-shear-layer reattachment point. However, the distributions began to diverge near the reattachment point, with the experimental heating becoming increasingly greater than the computed heating with distance downstream from the reattachment point. It was concluded that this divergence was due to transition of the wake free shear layer just upstream of the reattachment point on the sting.
Quantitative observations of transitional boundary layers in regions of strong flow deceleration on an axial compressor stator blade are reported. Measurements are obtained at a fixed chordwise position, and the blade incidence was varied by changing the compressor throughflow so as to move the transition region relative to the stationary probe. It was thus possible to observe typical boundary layer behavior at various stages of transition in the turbomachine environment. The range of observations covers separating laminar flow at transition onset, and reattachment of intermittently turbulent periodically separated shear layers.
Insights and lessons learned from the aerodynamic analysis of the High Wing Transport (HWT) high-lift configuration are presented. Three-dimensional Navier-Stokes CFD simulations using the OVERFLOW flow solver are compared with high Reynolds test data obtained in the NASA Ames 12 Foot Pressure Wind Tunnel (PWT) facility. Computational analysis of the baseline HWT high-lift configuration with and without Externally Blown Flap (EBF) jet effects is highlighted. Several additional aerodynamic investigations, such as nacelle strake effectiveness and wake vortex studies, are presented. Technical capabilities and shortcomings of the computational method are discussed and summarized.
Solar wind plasma effects on lunar wake umbral and penumbral interplanetary magnetic field anomalies, evaluating roles of electrons and ions
Measurements of the time-averaged induced velocities were obtained for rotor tip speeds as great as 1,100 feet per second (tip Mach number of 0.98) and measurements of the instantaneous induced velocities were obtained for rotor tip speeds as great as 900 feet per second. The results indicate that the small effects on the wake with increasing Mach number are primarily due to the changes in rotor-load distribution resulting from changes in Mach number rather than to compressibility effects on the wake itself. No effect of tip Mach number on the instantaneous velocities was observed. Under conditions for which the blade tip was operated at negative pitch angles, an erratic circulatory flow was observed.
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.
An experimental investigation of the effects of the speed of rotation on a turbomachinery rotor blade wake was conducted using a fan rotor in incompressible flow. Measurements were made at two different rotational speeds (1753 and 1010 rpm) but keeping the same blade loading. This was achieved by maintaining the same incidence angle to the rotor blades in both sets of measurements. The blade incidence angle was also varied to discern the effect of blade loading. A tri-axial hot wire probe mounted in a stationary frame of reference was used for the measurements made at several radial and axial stations in the near- and far-wake regions. The three dimensional mean velocity and turbulence profiles, the wake defect, wake decay rate, turbulence intensities and turbulence stresses are appreciably altered when the speed of rotation and the blade loading are changed. The wake defect is substantially reduced and the radial velocity increased when the rotational speed is increased.