ALBERT

Description: Wind tunnel experiments will continue to be a primary source of validation data for many types of mathematical and computational models in the aerospace industry. The increased emphasis on accuracy of data acquired from these facilities requires understanding of the uncertainty of not only the measurement data but also any correction applied to the data. One of the largest and most critical corrections made to these data is due to wall interference. In an effort to understand the accuracy and suitability of these corrections, a statistical validation process for wall interference correction methods has been developed. This process is based on the use of independent cases which, after correction, are expected to produce the same result. Comparison of these independent cases with respect to the uncertainty in the correction process establishes a domain of applicability based on the capability of the method to provide reasonable corrections with respect to customer accuracy requirements. The statistical validation method was applied to the version of the Transonic Wall Interference Correction System (TWICS) recently implemented in the National Transonic Facility at NASA Langley Research Center. The TWICS code generates corrections for solid and slotted wall interference in the model pitch plane based on boundary pressure measurements. Before validation could be performed on this method, it was necessary to calibrate the ventilated wall boundary condition parameters. Discrimination comparisons are used to determine the most representative of three linear boundary condition models which have historically been used to represent longitudinally slotted test section walls. Of the three linear boundary condition models implemented for ventilated walls, the general slotted wall model was the most representative of the data. The TWICS code using the calibrated general slotted wall model was found to be valid to within the process uncertainty for test section Mach numbers less than or equal to 0.60. The scatter among the mean corrected results of the bodies of revolution validation cases was within one count of drag on a typical transport aircraft configuration for Mach numbers at or below 0.80 and two counts of drag for Mach numbers at or below 0.90.

Description: An investigation of the aerothermodynamic environment of the Huygens entry probe has been conducted. A Monte Carlo simulation of the trajectory of the probe during entry into Titan's atmosphere was performed to identify a worst-case heating rate trajectory. Flowfield and radiation transport computations were performed at points along this trajectory to obtain convective and radiative heat-transfer distributions on the probe's heat shield. This investigation identified important physical and numerical factors, including atmospheric CH4 concentration, transition to turbulence, numerical diffusion modeling, and radiation modeling, which strongly influenced the aerothermodynamic environment.

Description: Computational Fluid Dynamics (CFD) is increasingly being used to both augment and create an aerodynamic performance database for aircraft configurations. This aerodynamic database contains the response of the aircraft to varying flight conditions and control surface deflections. The current work presents a novel method for calculating dynamic stability derivatives which reduces the computational cost over traditional unsteady CFD approaches by an order of magnitude, while still being applicable to arbitrarily complex geometries over a wide range of flow regimes. The primary thesis of this work is that the response to a forced motion can often be represented with a small, predictable number of frequency components without loss of accuracy. By resolving only those frequencies of interest, the computational effort is significantly reduced so that the routine calculation of dynamic derivatives becomes practical. The current implementation uses this same non-linear, frequency-domain approach and extends the application to the 3-D Euler equations. The current work uses a Cartesian, embedded-boundary method to automate the generation of dynamic stability derivatives.

Description: A capability for real-time computational simulation of aeroheating has been developed in support of the Hyper-X program, which is directed toward demonstrating the feasibility of operating an air-breathing ramjet/scramjet engine at mach 5, mach 7, and mach 10. The simulation software will serve as a valuable design tool for initial trajectory studies in which aerodynamic heating is expected to exert a major influence in the design of the Hyper-X airplane; this tool will aid in the selection of materials, sizing of structural skin thicknesses, and selection of components of a thermal-protection system (TPS) for structures that must be insulated against aeroheating.

Description: A rotor blade system with reduced blade-vortex interaction noise includes a plurality of tube members embedded in proximity to a tip of each rotor blade. The inlets of the tube members are arrayed at the leading edge of the blade slightly above the chord plane, while the outlets are arrayed at the blade tip face. Such a design rapidly diffuses the vorticity contained within the concentrated tip vortex because of enhanced flow mixing in the inner core, which prevents the development of a laminar core region.

Description: A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76 deg/40 deg double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 30 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M = 0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Description: An airfoil having a fore airfoil element, an aft airfoil element, and a slot region in between them. These elements induce laminar flow over substantially all of the fore airfoil element and also provide for laminar flow in at least a portion of the slot region. The method of the invention is one for inducing natural laminar flow over an airfoil. In the method, a fore airfoil element, having a leading and trailing edge, and an aft airfoil element define a slot region. Natural laminar flow is induced over substantially all of the fore airfoil element, by inducing the pressures on both surfaces of the fore airfoil element to decrease to a location proximate the trailing edge of the fore airfoil element using pressures created by the aft airfoil element.

Description: A multi-element rotor blade includes an individually controllable main element and fixed aerodynamic surface in an aerodynamically efficient location relative to the main element. The main element is controlled to locate the fixed aerodynamic surface in a position to increase lift and/or reduce drag upon the main element at various azimuthal positions during rotation.

Description: An experimental study has been performed to develop a large force and moment aerodynamic data set on a slender axisymmetric missile configuration having cruciform strakes and in-line control tail fins. The data include six-component balance measurements of the configuration aerodynamics and three-component measurements on all four tail fins. The test variables include angle of attack, roll angle, Mach number, model buildup, strake length, nose size, and tail fin deflection angles to provide pitch, yaw, and roll control. Test Mach numbers ranged from 0.60 to 4.63. The entire data set is presented on a CD-ROM that is attached to this paper. The CD-ROM also includes extensive plots of both the six-component configuration data and the three-component tail fin data. Selected samples of these plots are presented in this paper to illustrate the features of the data and to investigate the effects of the test variables.

Description: A wind tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing capability. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at momentum coefficients (Cm) from 0.0 to 0.12. Test data were acquired at Mach numbers of 0.3, 0.5, 0.7, 0.8, and 0.84 at Reynolds numbers per foot of 2.43 x 105 to 1.05 x 106. For a transonic condition, (Mach = 0.8 at alpha = 3 degrees), it was generally found the smaller slot and larger Coanda surface combination was overall more effective than other slot/Coanda surface combinations. Lower surface blowing was not as effective as the upper surface blowing over the same range of momentum coefficients. No appreciable Coanda surface, slot height, or slot blowing position preference was indicated transonically with the dual slot blowing.

Description: During Phase I of this project, Raytheon Aircraft Company (RAC) has analytically and experimentally evaluated key components of a system that could be implemented for active tailoring of wing lift distribution using low-drag, trailing-edge modifications. Simple systems such as those studied by RAC could be used to enhance the cruise performance of a business jet configuration over a range of typical flight conditions. The trailing-edge modifications focus on simple, deployable mechanisms comprised of extendable small flap panels over portions of the span that could be used to subtly but positively optimize the lift and drag characteristics. The report includes results from low speed wind tunnel testing of the trailing-edge devices, descriptions of potential mechanisms for automation, and an assessment of the technology.

Description: Method and system for analyzing aircraft data, including multiple selected flight parameters for a selected phase of a selected flight, and for determining when the selected phase of the selected flight is atypical, when compared with corresponding data for the same phase for other similar flights. A flight signature is computed using continuous-valued and discrete-valued flight parameters for the selected flight parameters and is optionally compared with a statistical distribution of other observed flight signatures, yielding atypicality scores for the same phase for other similar flights. A cluster analysis is optionally applied to the flight signatures to define an optimal collection of clusters. A level of atypicality for a selected flight is estimated, based upon an index associated with the cluster analysis.

Description: This paper describes model structures and parameter estimation algorithms suitable for the identification of unsteady aerodynamic models from input-output data. The model structures presented are state space models and include linear time-invariant (LTI) models and linear parameter-varying (LPV) models. They cover a wide range of local and parameter dependent identification problems arising in unsteady aerodynamics and nonlinear flight dynamics. We present a residue algorithm for estimating model parameters from data. The algorithm can incorporate apriori information and is described in detail. The algorithms are evaluated on the F-16XL wind-tunnel test data from NAS Langley Research Center. Results of numerical evaluation are presented. The paper concludes with a discussion major issues and directions for future work.

Description: Significant advances have been made to non-intrusive flow field diagnostics in the past decade. Camera based techniques are now capable of determining physical qualities such as surface deformation, surface pressure and temperature, flow velocities, and molecular species concentration. In each case, extracting the pertinent information from the large volume of acquired data requires powerful and efficient data visualization tools. The additional requirement for real time visualization is fueled by an increased emphasis on minimizing test time in expensive facilities. This paper will address a capability titled LiveView3D, which is the first step in the development phase of an in depth, real time data visualization and analysis tool for use in aerospace testing facilities.

Description: The pulsed flow emitted from a shrouded Hartmann-Sprenger tube was sampled with high-frequency pressure transducers and with laser particle imaging velocimetry, and found to consist of a train of vortices. Thrust and mass flow were also monitored using a thrust plate and orifice, respectively. The tube and shroud lengths were altered to give four different operating frequencies. From the data, the radius, velocity, and circulation of the vortex rings was obtained. Each frequency corresponded to a different length to diameter ratio of the pulse of air leaving the driver shroud. Two of the frequencies had length to diameter ratios below the formation number, and two above. The formation number is the value of length to diameter ratio below which the pulse converts to a vortex ring only, and above which the pulse becomes a vortex ring plus a trailing jet. A modified version of the slug model of vortex ring formation was used to compare the observations with calculated values. Because the flow exit area is an annulus, vorticity is shed at both the inner and outer edge of the jet. This results in a reduced circulation compared with the value calculated from slug theory accounting only for the outer edge. If the value of circulation obtained from laser particle imaging velocimetry is used in the slug model calculation of vortex ring velocity, the agreement is quite good. The vortex ring radius, which does not depend on the circulation, agrees well with predictions from the slug model.

Description: The program objectives are fully defined in the original proposal entitled Program of Research in Flight Dynamics in GW at NASA Langley Research Center, which was originated March 20, 1975, and in the renewals of the research program from January 1, 2003 to September 30, 2005. The program in its present form includes three major topics: 1. the improvement of existing methods and development of new methods for wind tunnel and flight data analysis, 2. the application of these methods to wind tunnel and flight test data obtained from advanced airplanes, 3. the correlation of flight results with wind tunnel measurements, and theoretical predictions.

Description: Computational aerothermodynamic simulations of Orbiter windside tile damage in flight were performed in support of the Space Shuttle Return-to-Flight effort. The simulations were performed for both hypervelocity flight and low-enthalpy wind tunnel conditions and contributed to the Return-to-Flight program by providing information to support a variety of damage scenario analyses. Computations at flight conditions were performed at or very near the peak heating trajectory point for multiple damage scenarios involving damage windside acreage reaction cured glass (RCG) coated silica tile(s). The cavities formed by the missing tile examined in this study were relatively short leading to flow features which indicated open cavity behavior. Results of the computations indicated elevated heating bump factor levels predicted for flight over the predictions for wind tunnel conditions. The peak heating bump factors, defined as the local heating to a reference value upstream of the cavity, on the cavity floor for flight simulation were 67% larger than the peak wind tunnel simulation value. On the downstream face of the cavity the flight simulation values were 60% larger than the wind tunnel simulation values. On the outer mold line (OML) downstream of the cavity, the flight values are about 20% larger than the wind tunnel simulation values. The higher heating bump factors observed in the flight simulations were due to the larger driving potential in terms of energy entering the cavity for the flight simulations. This is evidenced by the larger rate of increase in the total enthalpy through the boundary layer prior to the cavity for the flight simulation.

Description: An analysis of the flow state on a trapezoidal wing model from the NASA 3-D High Lift Flow Physics Experiment is presented. The objective of the experiment was to characterize the flow over a non-proprietary semi-span three-element high-lift configuration to aid in assessing the state of the art in the computation of three-dimensional high-lift flows. Surface pressures and hot-film sensors are used to determine the flow conditions on the slat, main, and flap. The locations of the attachments lines and the values of the attachment line Reynolds number are estimated based on the model surface pressures. Data from the hot-films are used to determine if the flow is laminar, transitional, or turbulent by examining the hot-film time histories, statistics, and frequency spectra.

Description: Boundary layer receptivity to two-dimensional slow and fast acoustic waves is investigated by solving Navier-Stokes equations for Mach 4.5 flow over a flat plate with a finite-thickness leading edge. Higher order spatial and temporal schemes are employed to obtain the solution whereby the flat-plate leading edge region is resolved by providing a sufficiently refined grid. The results show that the instability waves are generated in the leading edge region and that the boundary-layer is much more receptive to slow acoustic waves (by almost a factor of 20) as compared to the fast waves. Hence, this leading-edge receptivity mechanism is expected to be more relevant in the transition process for high Mach number flows. The effect of acoustic wave incidence angle is also studied and it is found that the receptivity of the boundary layer on the windward side (with respect to the acoustic forcing) decreases by more than a factor of 4 when the incidence angle is increased from 0 to 45 deg. However, the receptivity coefficient for the leeward side is found to vary relatively weakly with the incidence angle. The effect of leading-edge thickness is also studied and bluntness is found to stabilize the boundary layer. The relative significance of fast acoustic waves is enhanced in the presence of bluntness.

Description: The Lag model has shown great promise in prediction of low speed and transonic separations. The predictions of the model, along with other models (Spalart-Allmaras and Menter SST) are assessed for various high speed flowfields. In addition to skin friction and separation predictions, the prediction of heat transfer are compared among these models, and some fundamental building block flowfields, are investigated.

Description: The aerodynamic performance of an isolated fan or rotor alone model was measured in the NASA Glenn Research Center 9- by 15- Foot Low Speed Wind Tunnel as part of the Fan Broadband Source Diagnostic Test conducted at NASA Glenn. The Source Diagnostic Test was conducted to identify the noise sources within a wind tunnel scale model of a turbofan engine and quantify their contribution to the overall system noise level. The fan was part of a 1/5th scale model representation of the bypass stage of a current technology turbofan engine. For the rotor alone testing, the fan and nacelle, including the inlet, external cowl, and fixed area fan exit nozzle, were modeled in the test hardware; the internal outlet guide vanes located behind the fan were removed. Without the outlet guide vanes, the velocity at the nozzle exit changes significantly, thereby affecting the fan performance. As part of the investigation, variations in the fan nozzle area were tested in order to match as closely as possible the rotor alone performance with the fan performance obtained with the outlet guide vanes installed. The fan operating performance was determined using fixed pressure/temperature combination rakes and the corrected weight flow. The performance results indicate that a suitable nozzle exit was achieved to be able to closely match the rotor alone and fan/outlet guide vane configuration performance on the sea level operating line. A small shift in the slope of the sea level operating line was measured, which resulted in a slightly higher rotor alone fan pressure ratio at take-off conditions, matched fan performance at cutback conditions, and a slightly lower rotor alone fan pressure ratio at approach conditions. However, the small differences in fan performance at all fan conditions were considered too small to affect the fan acoustic performance.

Description: The goal of the DARPA Shaped Sonic Boom Demonstration (SSBD) Program was to demonstrate for the first time in flight that sonic booms can be substantially reduced by incorporating specialized aircraft shaping techniques. Although mitigation of the sonic boom via specialized shaping techniques was theorized decades ago, until now, this theory had never been tested with a flight vehicle subjected to actual flight conditions in a real atmosphere. The demonstrative success, which occurred on 27 August 2003 with repeat flights in the supersonic corridor at Edwards Air Force Base, is a critical milestone in the development of next generation supersonic aircraft that could one day fly unrestricted over land and help usher in a new era of time-critical air transport. Pressure measurements obtained on the ground and in the air confirmed that the specific modifications made to a Northrop Grumman F-5E aircraft not only changed the shape of the shock wave signature emanating from the aircraft, but also produced a flat-top signature whose shape persisted, as predicted, as the pressure waves propagated through the atmosphere to the ground. This accomplishment represents a major advance towards reducing the startling and potentially damaging noise of a sonic boom. This paper describes the evolution of the SSBD program, including the rationale for test article selection, and provides an overview of the history making accomplishments achieved during the SSBD effort, as well as, the follow-on NASA Shaped Sonic Boom Experiment (SSBE) Program, whose goal was to further evaluate the characteristics and robustness of shaped boom signatures.

Description: Turbulent separated flow over a two-dimensional hump is computed by solving the RANS equations with k - omega (SST) turbulence model for the baseline, steady suction and oscillatory blowing/suction flow control cases. The flow equations and the turbulent model equations are solved using a fifth-order accurate weighted essentially. nonoscillatory (WENO) scheme for space discretization and a third order, total variation diminishing (TVD) Runge-Kutta scheme for time integration. Qualitatively the computed pressure distributions exhibit the same behavior as those observed in the experiments. The computed separation regions are much longer than those observed experimentally. However, the percentage reduction in the separation region in the steady suction case is closer to what was measured in the experiment. The computations did not predict the expected reduction in the separation length in the oscillatory case. The predicted turbulent quantities are two to three times smaller than the measured values pointing towards the deficiencies in the existing turbulent models when they are applied to strong steady/unsteady separated flows.

Description: The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.

Description: This viewgraph presentation reviews NASA's basics of flight, controlling flight, and flight systems.

Description: Turbulent thin-layer, Reynolds-Averaged Navier-Stokes solutions, based on a multi-block structured grid, are presented for a 65 deg delta wing having either a sharp leading edge (SLE) or blunt leading edge (BLE) geometry. The primary objective of the study is to assess the prediction capability of the method for simulating the leading-edge flow separation and the ensuing vortex flow characteristics. Computational results are obtained for two angles of attack of approximately 13 and 20 deg, at free-stream Mach number of 0.40 and Reynolds number of 6 million based on the wing mean aerodynamic chord. The effects of two turbulence models of Baldwin-Lomax with Degani-Schiff (BL/DS) and the Spalart-Allmaras (SA) on the numerical results are also discussed. The computations also explore the effects of two numerical flux-splitting schemes, i.e., flux difference splitting (fds) and flux vector splitting (fvs), on the solution development and convergence characteristics. The resulting trends in solution sensitivity to grid resolution for the selected leading-edge geometries, angles of attack, turbulence models and flux splitting schemes are also presented. The validity of the numerical results is evaluated against a unique set of experimental wind-tunnel data that was obtained in the National Transonic Facility at the NASA Langley Research Center.

Description: An experimental study has been performed to develop a large force and moment aerodynamic data set on a slender axisymmetric missile configuration having cruciform strakes and in-line control tail fins. The data include six-component balance measurements of the configuration aerodynamics and three-component measurements on all four tail fins. The test variables include angle of attack, roll angle, Mach number, model buildup, strake length, nose size, and tail fin deflection angles to provide pitch, yaw, and roll control. Test Mach numbers ranged from 0.60 to 4.63. The entire data set is presented on a CD-ROM that is attached to this paper. The CD-ROM also includes extensive plots of both the six-component configuration data and the three-component tail fin data. Selected samples of these plots are presented in this paper to illustrate the features of the data and to investigate the effects of the test variables.

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 Path finder, 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: Effects of buoyancy on transition from laminar to turbulent flow are presented for momentum-dominated helium jet injected into ambient air. The buoyancy was varied in a 2.2-sec drop tower facility without affecting the remaining operating parameters. The jet flow in Earth gravity and microgravity was visualized using the rainbow schlieren deflectometry apparatus. Results show significant changes in the flow structure and transition behavior in the absence of buoyancy.

Description: A dynamic wake avoidance system utilizes aircraft and atmospheric parameters readily available in flight to model and predict airborne wake vortices in real time. A novel combination of algorithms allows for a relatively simple yet robust wake model to be constructed based on information extracted from a broadcast. The system predicts the location and movement of the wake based on the nominal wake model and correspondingly performs an uncertainty analysis on the wake model to determine a wake hazard zone (no fly zone), which comprises a plurality of wake planes, each moving independently from another. The system selectively adjusts dimensions of each wake plane to minimize spatial and temporal uncertainty, thereby ensuring that the actual wake is within the wake hazard zone. The predicted wake hazard zone is communicated in real time directly to a user via a realistic visual representation. In an example, the wake hazard zone is visualized on a 3-D flight deck display to enable a pilot to visualize or see a neighboring aircraft as well as its wake. The system substantially enhances the pilot's situational awareness and allows for a further safe decrease in spacing, which could alleviate airport and airspace congestion.

Description: A computational study of the separated flow through a 2-D asymmetric subsonic diffuser has been performed. The Wind Computational Fluid Dynamics code is used to predict the separation and reattachment behavior for an incompressible diffuser flow. The diffuser inlet flow is a two-dimensional, turbulent, and fully-developed channel flow with a Reynolds number of 20,000 based on the centerline velocity and the channel height. Wind solutions computed with the Menter SST, Chien k-epsilon, Spalart-Allmaras and Explicit Algebraic Reynolds Stress turbulence models are compared with experimentally measured velocity profiles and skin friction along the upper and lower walls. In addition to the turbulence model study, the effects of grid resolution and use of wall functions were investigated. The grid studies varied the number of grid points across the diffuser and varied the initial wall spacing from y(sup +) = 0.2 to 60. The wall function study assessed the applicability of wall functions for analysis of separated flow. The SST and Explicit Algebraic Stress models provide the best agreement with experimental data, and it is recommended wall functions should only be used with a high level of caution.

Description: A method for estimating the sonic-boom overpressures from a conceptual aircraft where the lift is carried by both a canard and a wing during supersonic cruise is presented and discussed. Computer codes used for the prediction of the aerodynamic performance of the wing, the canard-wing interference, the nacelle-wing interference, and the sonic-boom overpressures are identified and discussed as the procedures in the method are discussed. A canard-wing supersonic-cruise concept was used as an example to demonstrate the application of the method.

Description: An investigation using a survey rake with 11 five-hole pyramid-head probes has been conducted in the Langley Transonic Dynamics Tunnel (TDT) to measure the test section flow angularity. Flow measurements were made in a 10-ft square grid centered about the test section centerline at a single streamwise location for nine Mach numbers ranging from 0.50 to 1.19 at dynamic pressures of 100 and 225 pounds per square foot. Test section flow angularity was found to be minimal with a generally random flow pattern. Corrections for survey rake induced in-plane flow were determined to be necessary; however, corrections for rake induced lift effects were not required.

Description: A new computer program has been developed called ASP3D (Advanced Small Perturbation 3D), which solves the small perturbation potential flow equation in an advanced form including mass-consistent surface and trailing wake boundary conditions, and entropy, vorticity, and viscous effects. The purpose of the program is for unsteady aerodynamic and aeroelastic analyses, especially in the nonlinear transonic flight regime. The program exploits the simplicity of stationary Cartesian meshes with the movement or deformation of the configuration under consideration incorporated into the solution algorithm through a planar surface boundary condition. The new ASP3D code is the result of a decade of developmental work on improvements to the small perturbation formulation, performed while the author was employed as a Senior Research Scientist in the Configuration Aerodynamics Branch at the NASA Langley Research Center. The ASP3D code is a significant improvement to the state-of-the-art for transonic aeroelastic analyses over the CAP-TSD code (Computational Aeroelasticity Program Transonic Small Disturbance), which was developed principally by the author in the mid-1980s. The author is in a unique position as the developer of both computer programs to compare, contrast, and ultimately make conclusions regarding the underlying formulations and utility of each code. The paper describes the salient features of the ASP3D code including the rationale for improvements in comparison with CAP-TSD. Numerous results are presented to demonstrate the ASP3D capability. The general conclusion is that the new ASP3D capability is superior to the older CAP-TSD code because of the myriad improvements developed and incorporated.

Description: The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

Description: The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the physics of the separation phenomenon under periodic unsteady wake flow. Several physical mechanisms are discussed.

Description: A method for reducing drag upon a blunt-based vehicle by adaptively increasing forebody roughness to increase drag at the roughened area of the forebody, which results in a decrease in drag at the base of this vehicle, and in total vehicle drag.

Description: A high Reynolds number wind tunnel test was conducted to assess Reynolds number effects on the aerodynamic performance characteristics of a realistic, second-generation supersonic transport concept. The tests included longitudinal studies at transonic and low-speed, high-lift conditions across a range of chord Reynolds numbers (8 million to 120 million). Results presented focus on Reynolds number and static aeroelastic sensitivities at Mach 0.30 and 0.90 for a configuration without a tail. Static aeroelastic effects, which mask Reynolds number effects, were observed. Reynolds number effects were generally small and the drag data followed established trends of skin friction as a function of Reynolds number. A more nose-down pitching moment was produced as Reynolds number increased because of an outward movement of the inboard leading-edge separation at constant angles of attack. This study extends the existing Reynolds number database for supersonic transports operating at off-design conditions.

Description: The influence of vortex ring state (VRS) on rotorcraft flight dynamics is investigated, specifically the vertical velocity drop of helicopters and the roll-off of tiltrotors encountering VRS. The available wind tunnel and flight test data for rotors in vortex ring state are reviewed. Test data for axial flow, non-axial flow, two rotors, unsteadiness, and vortex ring state boundaries are described and discussed. Based on the available measured data, a VRS model is developed. The VRS model is a parametric extension of momentum theory for calculation of the mean inflow of a rotor, hence suitable for simple calculations and real-time simulations. This inflow model is primarily defined in terms of the stability boundary of the aircraft motion. Calculations of helicopter response during VRS encounter were performed, and good correlation is shown with the vertical velocity drop measured in flight tests. Calculations of tiltrotor response during VRS encounter were performed, showing the roll-off behavior characteristic of tiltrotors. Hence it is possible, using a model of the mean inflow of an isolated rotor, to explain the basic behavior of both helicopters and tiltrotors in vortex ring state.

Description: An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.This CD-ROM supplement contains inlet data including: Boundary layer data, Duct static pressure data, performance-AIP (fan face) data, Photos, Tunnel wall P-PTO data and definitions.

Description: Wind tunnel experiments will continue to be a primary source of validation data for many types of mathematical and computational models in the aerospace industry. The increased emphasis on accuracy of data acquired from these facilities requires understanding of the uncertainty of not only the measurement data but also any correction applied to the data. One of the largest and most critical corrections made to these data is due to wall interference. In an effort to understand the accuracy and suitability of these corrections, a statistical validation process for wall interference correction methods has been developed. This process is based on the use of independent cases which, after correction, are expected to produce the same result. Comparison of these independent cases with respect to the uncertainty in the correction process establishes a domain of applicability based on the capability of the method to provide reasonable corrections with respect to customer accuracy requirements. The statistical validation method was applied to the version of the Transonic Wall Interference Correction System (TWICS) recently implemented in the National Transonic Facility at NASA Langley Research Center. The TWICS code generates corrections for solid and slotted wall interference in the model pitch plane based on boundary pressure measurements. Before validation could be performed on this method, it was necessary to calibrate the ventilated wall boundary condition parameters. Discrimination comparisons are used to determine the most representative of three linear boundary condition models which have historically been used to represent longitudinally slotted test section walls. Of the three linear boundary condition models implemented for ventilated walls, the general slotted wall model was the most representative of the data. The TWICS code using the calibrated general slotted wall model was found to be valid to within the process uncertainty for test section Mach numbers less than or equal to 0.60. The scatter among the mean corrected results of the bodies of revolution validation cases was within one count of drag on a typical transport aircraft configuration for Mach numbers at or below 0.80 and two counts of drag for Mach numbers at or below 0.90.

Description: An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.

Description: A free-to-roll study of the low-speed lateral characteristics of the pre-production F/A-18E was conducted in the NASA Langley 12-Foot Low-Speed Tunnel. In developmental flight tests the F/A-18E unexpectedly experienced uncommanded lateral motions in the power approach configuration. The objective of this study was to determine the feasibility of using the free-to-roll technique for the detection of uncommanded lateral motions for the pre-production F/A-18E in the power approach configuration. The data revealed that this technique in conjunction with static data revealed insight into the cause of the lateral motions. The free-to-roll technique identified uncommanded lateral motions at the same angle-of-attack range as experienced in flight tests. The cause of the uncommanded lateral motions was unsteady asymmetric wing stall. The paper also shows that free-to-roll data or static force and moment data alone are not enough to accurately capture the potential for an aircraft to experience uncommanded lateral motion.

Description: A new data analysis technique for the identification of static and dynamic aerodynamic stability coefficients from wind tunnel test video data is presented. This new technique was applied to video data obtained during a parachute wind tunnel test program conducted in support of the Mars Exploration Rover Mission. Total angle-of-attack data obtained from video images were used to determine the static pitching moment curve of the parachute. During the original wind tunnel test program the static pitching moment curve had been determined by forcing the parachute to a specific total angle-of -attack and measuring the forces generated. It is shown with the new technique that this parachute, when free to rotate, trims at an angle-of-attack two degrees lower than was measured during the forced-angle tests. An attempt was also made to extract pitch damping information from the video data. Results suggest that the parachute is dynamically unstable at the static trim point and tends to become dynamically stable away from the trim point. These trends are in agreement with limit-cycle-like behavior observed in the video. However, the chaotic motion of the parachute produced results with large uncertainty bands.

Description: A pilot study was conducted on a flapped semi-span model to investigate the concept and viability of near-wake vortex management via separation control. Passive control was achieved by means of a simple fairing and active control was achieved via zero mass-flux blowing slots. Vortex sheet strength, estimated by integrating surface pressure ports, was used to predict vortex characteristics by means of inviscid rollup relations. Furthermore, vortices trailing the flaps were mapped using a seven-hole probe. Separation control was found to have a marked effect on vortex location, strength, tangential velocity, axial velocity and size over a wide range of angles of attack and control conditions. In general, the vortex trends were well predicted by the inviscid rollup relations. Manipulation of the separated flow near the flap edges exerted significant control over both outboard and inboard edge vortices while producing negligible lift excursions. Dynamic separation and attachment control was found to be an effective means for dynamically perturbing the vortex from arbitrarily long wavelengths down to wavelengths less than a typical wingspan. In summary, separation control has the potential for application to time-independent or time-dependent wake alleviation schemes, where the latter can be deployed to minimize adverse effects on ride-quality and dynamic structural loading.

Description: An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface.

Description: In October 2003, NASA embarked on the ACAST project (Advanced CNS Architectures and System Technologies) to perform research and development on selected communications, navigation, and surveillance (CNS) technologies to enhance the performance of the National Airspace System (NAS). The Networking Research Group of NASA's ACAST project, in order to ensure global interoperability and deployment, formulated their own salient list of requirements. Many of these are not necessarily of concern to the FAA, but are a concern to those who have to deploy, operate, and pay for these systems. These requirements were submitted to the world s industries, governments, and academic institutions for comments. The results of that request for comments are summarized in this paper.

Description: Two multi-scale-type turbulence models are implemented in the PAB3D solver. The models are based on modifying the Reynolds Averaged Navier-Stokes (RANS) equations. The first scheme is a hybrid RANS/LES model utilizing the two-equation (k(epsilon)) model with a RANS/LES transition function dependent on grid spacing and the computed turbulence length scale. The second scheme is a modified version of the Partially Averaged Navier-Stokes (PANS) model, where the unresolved kinetic energy parameter (f(sub k)) is allowed to vary as a function of grid spacing and the turbulence length scale. This parameter is estimated based on a novel two-stage procedure to efficiently estimate the level of scale resolution possible for a given flow on a given grid for Partial Averaged Navier-Stokes (PANS). It has been found that the prescribed scale resolution can play a major role in obtaining accurate flow solutions. The parameter f(sub k) varies between zero and one and equal to one in the viscous sub layer, and when the RANS turbulent viscosity becomes smaller than the LES viscosity. The formulation, usage methodology, and validation examples are presented to demonstrate the enhancement of PAB3D's time-accurate and turbulence modeling capabilities. The accurate simulations of flow and turbulent quantities will provide valuable tool for accurate jet noise predictions. Solutions from these models are compared to RANS results and experimental data for high-temperature jet flows. The current results show promise for the capability of hybrid RANS/LES and PANS in simulating such flow phenomena.

Description: This paper presents the results of direct drag measurements on a variety of porous plate acoustic liners. The existing literature describes numerous studies of drag on porous walls with injection or suction, but relatively few of drag on porous plates with neither injection nor suction. Furthermore, the porosity of the porous plate in existing studies is much lower than typically used in acoustic liners. In the present work, the acoustic liners consisted of a perforated face sheet covering a bulk acoustic absorber material. Factors that were varied in the experiment were hole diameter, hole pattern, face sheet thickness, bulk material type, and size of the gap (if any) between the face sheet and the absorber material.

Description: As part of the research effort at NASA in support of the stage separation and ascent aerothermodynamics research program, proximity testing of a generic bimese wing-body configuration was conducted in NASA Langley's Aerothermodynamics Laboratory in the 20-Inch Mach 6 Air Tunnel. The objective of this work is the development of experimental tools and testing methodologies to apply to hypersonic stage separation problems for future multi-stage launch vehicle systems. Aerodynamic force and moment proximity data were generated at a nominal Mach number of 6 over a small range of angles of attack. The generic bimese configuration was tested in a belly-to-belly and back-to-belly orientation at 86 relative proximity locations. Over 800 aerodynamic proximity data points were taken to serve as a database for code validation. Longitudinal aerodynamic data generated in this test program show very good agreement with viscous computational predictions. Thus a framework has been established to study separation problems in the hypersonic regime using coordinated experimental and computational tools.

Description: Blade flap and chord bending and torsion moments are investigated for six rotors operating at transition and high speed: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale, and BO- 105 model (HART-I). The measured data from flight and wind tunnel tests are compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free wake models: rolled-up and multiple-trailer with consolidation models. At transition speed, there is fair to good agreement for the flap and chord bending moments between the test data and analysis for the H-34, research Puma, and SA 349/2. Torsion moment correlation, in general, is fair to good for all the rotors investigated. Better flap bending and torsion moment correlation is obtained for the UH-60A and BO-105 rotors by using the multiple-trailer with consolidation wake model. In the high speed condition, the analysis shows generally better correlation in magnitude than in phase for the flap bending and torsion moments. However, a significant underprediction of chord bending moment is observed for the research Puma and UH-60A. The poor chord bending moment correlation appears to be caused by the airloads model, not the structural dynamics.

Description: A new species of the genus Artocella van Achterberg is described from Spain. On the basis of its substantial sexual dimorphisim, the differences between the two previously known species of this Turanian-Mediterranean genus, which had each been described from specimens of only one (and differing) sex, are reassessed.

Description: Dasypoda intermedia spec. nov. from Iran is described. Its description fills a gap of our knowledge of the East Mediterranean fauna of the genus Dasypoda. The West Mediterranean Dasypoda species are well known but the eastern species lack convincing records. Moreover, D. intermedia spec. nov. is a very interesting species from a phylogenetic point of view. It shares some characters common to subgenera Dasypoda s. str. and Megadasypoda Michez, 2004, which provide further evidence for the close relationship of both subgenera.

Description: Thirty-one species of the family Platystictidae of the Philippines are revised, i.e. all species recognised, excluding the species of the Drepanosticta halterata-group. The following new taxa are described: 16 species in Drepanosticta Laidlaw: D. acuta spec. nov., D. aurita spec. nov., D. centrosaurus spec. nov., D. clados spec. nov., D. flavomaculata spec. nov., D. furcata spec. nov., D. hermes spec. nov., D. krios spec. nov., D. luzonica spec. nov., D. malleus spec. nov., D. myzouris spec. nov., D. paruatia spec. nov., D. pistor spec. nov., D. quadricornu spec. nov., D. rhamphis spec. nov., D. trachelocele spec. nov., two in Protosticta Selys, viz. P. lepteca spec. nov. and P. plicata spec. nov., and three in Sulcosticta gen. nov., viz. S. striata spec. nov., S. pallida spec. nov. and S. viticula spec. nov. The status of eleven previously described nominal taxa is established. One, D. septima Needham & Gyger, is doubtfully considered a synonym of D. mylitta Cowley. Based on a preliminary phylogenetic analysis, the species of Drepanosticta are divided into informal species groups. Most species of the Philippines have affinities to species of Sulawesi, the Moluccas and New Guinea. Several species confined to Palawan have sister-group relationships with species from Borneo. The affinities of various other species confined to the Sulu archipelago, are unsettled as yet. The species of Platystictidae here assigned to Protosticta Selys are presumably not closely related to the type species, P. simplicinervis Selys from Sulawesi. However, a better placement has to await a more detailed phylogenetic study of the family. For three species the new genus Sulcosticta gen. nov. is erected. These species are closely allied based on the structure of the appendages, but should have been assigned to different genera if based on the present generic definitions. Many species here described have small distributional ranges, a common phenomenon in Platystictidae. Since most forests in the Philippines are heavily under threat or have already disappeared in the last fifty years, several taxa described in this paper should be considered under threat of immediate extinction.

Description: In this part the remaining 78 species of the genus Pepsis, belonging to ten species-groups, are described and figured, and their phylogenetics and biogeography are discussed. 14 of the species are described as new: P. achterbergi spec. nov., P. adonta spec. nov., P. boharti spec. nov., P. caliente spec. nov., P. dayi spec. nov., P. esmeralda spec. nov., P. ianthoides spec. nov., P. jamaicensis spec. nov., P. krombeini spec. nov., P. martini spec. nov., P. multichroma spec. nov., P. nanoides spec. nov., P. wahisi spec. nov., and P. willinki spec. nov. Three species-names, P. infuscata Spinola, 1841, P. lampas Lucas, 1895, and P. thoreyi Dahlbom, 1845, are recalled from synonymy. The following 293 names are newly synonymized (the valid names are listed first): P. atalanta Mocsáry, 1885 = P. nitens Mocsáry, 1894, P. mocsaryi Lucas, 1895; P. inclyta Lepeletier, 1845 = P. mutabilis Lepeletier, 1845, P. vagabunda Lepeletier, 1845, P. cupripennis Taschenberg, 1869, P. violaceipennis Mocsáry, 1885, P. clotho Mocsáry, 1888, P. spengeli Mocsáry, 1888, P. sickmanni Mocsáry, 1888, P. nireus Mocsáry, 1894, P. atrovirens Lucas, 1895, P. cerastes Lucas, 1895, P. pygidialis Brèthes, 1908, P. guaranitica Brèthes, 1908, P. parca Lucas, 1919, P. atahualpa Banks, 1946, opimicornis, Haupt, 1952, atropos, Haupt, 1952, azurea Haupt, 1952; crassicornis Mocsáry, 1885 = P. sappho Brèthes, 1908, P. nitocris Brèthes, 1908, P. vivida Brèthes, 1908, P. arechavaletai Brèthes, 1908, P. lynchii Brèthes, 1908, P. operosa Brèthes, 1908, P. ataraqua Banks, 1946, P. splendida Haupt, 1952; P. sommeri Dahlbom, 1845 = P. azteca Cameron, 1893; P. xanthocera Dahlbom, 1843 = P. nigrescens Smith, 1855, P. fulgidipennis Mocsáry, 1885, P. juno Brèthes, 1908, P. ismare Banks, 1946, P. nigroprasina Haupt, 1952; P. seifferti Lucas, 1895 = P. cornuta Lucas, 1895, P. moebiusi Lucas, 1895, P. stygia Lucas, 1895; P. luteicornis Fabricius, 1804 = P. strenua Erichson, 1848, P. tinctipennis Smith, 1873, P. citreicornis Mocsáry, 1894, P. venosa Banks, 1945, P. alector Banks, 1946; P. asteria Mocsáry, 1894 = P. luridicornis Brèthes, 1926; P. convexa Lucas, 1895 = P. humeralis Brèthes, 1914; P. helvolicornis Lucas, 1895 = P. bahiae Brèthes, 1914; P. vitripennis Smith, 1855 = P. obscura Lepeletier, 1845, P. amabilis Mocsáry, 1885, P. centralis Cameron, 1893, P. margarete Lucas, 1895, P. venezuelae Kaye, 1913, P. aeneipennis Banks, 1946, P. helenae Haupt, 1952, P. coeruleoviridis Haupt, 1952; P. fumipennis Smith, 1855 = P. pallidicornis Mocsáry, 1885; P. amyntas Mocsáry, 1885 = P. vicina Lucas, 1895, P. clarinervis Brèthes, 1908, P. amyntoides Lucas, 1919, P. eurydice Lucas, 1919; P. dimidiata Fabricius, 1804 = P. vittigera Lucas, 1897, P. argentina Brèthes, 1908, P. sanctaeannae Brèthes, 1908, P. virgo Brèthes, 1908, P. externa Brèthes, 1908, P. transversa Brèthes, 1908, P. cordubensis Brèthes, 1908, P. banghaasi Lucas, 1919; P. menechma Lepeletier, 1845 = P. elegans Lepeletier, 1845, P. dubitata Cresson, 1867, P. prismatica Smith, 1855, P. advena Mocsáry, 1885, cinctipennis Mocsáry, 1885, P. guatemalensis Cameron, 1893, P. nestor Mocsáry, 1894, P. nigricornis Mocsáry, 1894, P. auranticornis Lucas, 1895, P. fruhstorferi Lucas, 1895, P. concolor Lucas, 1895, P. cerberus Lucas, 1895, P. euchroma Lucas, 1895, P. nigrocincta Lucas, 1895, P. mordax Lucas, 1895, P. inermis Fox, 1898, P. roberti Brèthes, 1908, P. janira Brèthes, 1908, P. cultrata Brèthes, 1908, P. novitia Banks, 1921; P. decipiens Lucas, 1895 = P. similis Lucas, 1895; P. minarum Brèthes, 1914 = P. pulchra Brèthes, 1914; P. basifusca Lucas, 1895 = P. angustimarginata Viereck, 1908; P. chrysoptera Burmeister, 1872 = P. exigua Lucas, 1895, P. smaragdinula Lucas, 1895, P. nebulosa Lucas, 1895, P. karschi Lucas, 1895, P. anisitsii Brèthes, 1908, P. indistincta Brèthes, 1908, P. dimidiatipennis Brèthes, 1908, P. chloroptera Brèthes, 1908, P. culta Brèthes, 1908, P. recta Brèthes, 1908, P. tornowii Brèthes, 1908, P. schrottkyi Brèthes, 1908, P. itinerata Brèthes, 1908, P. miniata Brèthes, 1908, P. spegazzinii Brèthes, 1908, P. paulistana Brèthes, 1914, P. chloe Brèthes, 1914, P. coronaria Brèthes, 1914, P. semilucana Haupt, 1952, P. bruneipes Haupt, 1952, P. brachynotus Haupt, 1952, P. diagonalis Haupt, 1952, P. discrepans Haupt, 1952; P. elongata Lepeletier, 1845 = P. purpurascens Smith, 1855, P. fuscipennis Smith, 1873, P. longula Banks, 1946; P. australis Saussure, 1867 = P. centaurus Lucas, 1897; P. cyanescens Lepeletier, 1845 = P. micans Mocsáry, 1885, P. jucunda Mocsáry, 1885, P. balloui Banks, 1946, P. diversa Haupt, 1952; P. lampas Lucas, 1895 = P. venturii Schrottky, 1902; P. nitida Lepeletier, 1845 = P. lucidula Smith, 1855, P. vaualba Smith, 1855, P. pruinosa Mocsáry, 1894, P. cylindrica Lucas, 1895, P. andina Brèthes, 1908, P. dilatata Brèthes, 1908, P. holmbergi Brèthes, 1908, P. concava Brèthes, 1908, P. ephebus Brèthes, 1908, P. vaga Brèthes, 1908, P. fuscobasalis Brèthes, 1908, P. cordata Brèthes, 1914, P. impatiens Brèthes, 1914, P. tricolor Brèthes, 1914, P. joergenseni Brèthes, 1914, P. cleone Brèthes, 1914, P. dorsata Brèthes, 1914, P. aretheas Brèthes, 1914, P. lassonis Lucas, 1819, P. consors Banks, 1946, P. interrupta Banks, 1946, P. analis Haupt, 1952; P. seladonica Dahlbom, 1843 = P. deuteroleuca Smith, 1855, P. kohli Lucas, 1895, P. venezolana Brèthes, 1908, P. burmeisteri Brèthes, 1908; P. cybele Banks, 1945 = P. weberi Banks, 1946; P. thoreyi Dahlbom, 1845 = P. lurida Lucas, 1895, P. euterpe Brèthes, 1908; P. flavescens Lucas, 1895 = P. periphetes Lucas, 1895, P. limbatella Brèthes, 1908, P. discoidalis Brèthes, 1914, P. limbatica Brèthes, 1914, P. militaris Brèthes, 1914, P. cavillatrix Haupt, 1952, P. arcuata Haupt, 1952, P. recterugosa Haupt, 1952, P. adversatrix Haupt, 1952; P. nigricans Lucas, 1895 = P. troglodytes Brèthes, 1908; P. montezuma Smith, 1855 = P. quitonensis Packard, 1869, P. sibylla Mocsáry, 1885, P. circe Mocsáry, 1885, P. occidentalis Cameron, 1893, P. peruanus Lucas, 1895, P. fulva Lucas, 1895, P. nessus Lucas, 1895, P. fusca Lucas, 1895, P. andicola Cameron, 1903, P. chilloensis Cameron, 1903, P. patagonica Brèthes, 1908, P. fasciculata Brèthes, 1908, P. pisoensis Strand, 1911, P. pacifica Brèthes, 1914, P. huascar Banks, 1946; P. completa Smith, 1855 = P. quichua Brèthes, 1908, P. comes Banks, 1946; P. smaragdina Dahlbom, 1843 = P. thunbergi Dahlbom, 1843, P. lara Mocsáry, 1888, P. satrapes Lucas, 1895, P. nupta Lucas, 1895, P. erynnis Lucas, 1895, P. fraterna Lucas, 1895, P. diabolus Lucas, 1895, P. mystica Lucas, 1895, P. thalia Brèthes, 1908, P. brasiliensis Brèthes, 1908, P. pallida Brèthes, 1908, P. iheringi Brèthes, 1908, P. dromeda Brèthes, 1908, P. sepultrix Lucas, 1919, P. strickeri Lucas, 1919; P. discolor Taschenberg, 1869 = P. sinnis Lucas, 1895, P. jujuyensis Brèthes, 1908, P. modesta Brèthes, 1908, P. comparata Brèthes, 1908, P. neutra Brèthes, 1908, P. terebrans Brèthes, 1908, P. procera Haupt, 1952, P. plaumanni Haupt, 1952, P. ogloblini Haupt, 1952, P. deletrix Haupt, 1952; P. limbata Guérin, 1831 = P. richteri Brèthes, 1908, P. polita Brèthes, 1908, P. limbella Haupt, 1952, P. artemis Haupt, 1952; P. basalis Mocsáry, 1885 = P. erdmanni Lucas, 1895, P. basinigra Haupt, 1952; P. infuscata Spinola, 1841 = P. niobe Mocsáry, 1885, P. sagana Mocsáry, 1894, P. incerta Banks, 1946; P. hyalinipennis Mocsáry, 1885 = P. subruficornis Haupt, 1952; P. festiva Fabricius, 1804 = P. pulchella Lepeletier, 1845, P. solitaria Smith, 1879, P. gallardoi Brèthes, 1908, P. hora Brèthes, 1914, P. amok Lucas, 1919, P. riojaneirensis Lucas, 1919; P. gracilis Lepeletier, 1845 = P. diana Mocsáry, 1885, P. hecate Mocsáry, 1885, P. spathulifera Lucas, 1895, P. sphinx Lucas, 1895, P. ierensis Banks, 1945, P. alceste Banks, 1946, P. scalaris Haupt, 1952; P. mildei Stål, 1857 = P. charon Mocsáry, 1885, P. cyanoptera Lucas, 1895, P. dryas Lucas, 1919; P. filiola Brèthes, 1914 = P. denserugosa Haupt, 1952; P. ruficornis Fabricius, 1804 = P. saphirus Palisot de Beauvois, 1805, P. violacea Mocsáry, 1885, P. hexamita Lucas, 1895, P. omniviolacea Haupt, 1952; P. brunneicornis Lucas, 1895 = P. glabripennis Lucas, 1895; P. purpurea Smith, 1873 = P. pan Mocsáry, 1885, P. parthenope Mocsáry, 1885, P. sagax Lucas, 1895, P. clypeata Brèthes, 1914, P. consimilis Banks, 1946, P. laconia Banks, 1946; P. viridisetosa Spinola, 1841 = P. eximia Smith, 1873; P. viridis Lepeletier, 1845 = P. errans Lepeletier, 1845, P. chlorotica Mocsáry, 1885, P. excelsa Lucas, 1895, P. selene Lucas, 1895, P. fimbriata Lucas, 1895, P. calypso Brèthes, 1908, P. fluminensis Brèthes, 1908, P. argentinicus Strand, 1910, P. mimetica Brèthes, 1914, P. garbei Brèthes, 1914, P. erecta Brèthes, 1914, P. tandilensis Brèthes, 1914, P. meridionalis Brèthes, 1914, P. minor Lucas, 1919, P. basifulgens Lucas, 1919, P. nebulosipennis Lucas, 1919, P. purpurea Lucas, 1919, P. koerberi Lucas, 1919, P. inimicissima Lucas, 1919, P. debilitans Lucas, 1919, P. itapaca Banks, 1946; P. aciculata Taschenberg, 1869 = P. nero Lucas, 1895; P. atripennis Fabricius, 1804 = P. flavilis Brèthes, 1908; P. ianthina Erichson, 1848 = P. fulvicornis Mocsáry, 1885, P. sirene Lucas, 1895, P. balboae Lucas, 1919, P. herodes Lucas, 1919, P. curti Lucas, 1919; P. nana Mocsáry, 1885 = P. mapiriensis Lucas, 1919, P. vinciens Lucas, 1919, P. ilione Banks, 1946, P. moesta Banks, 1946, P. orestes Banks, 1946, P. amautas Banks, 1946, P. inaequalis Haupt, 1952; P. hirtiventris Banks, 1946 = P. viridaurea Haupt, 1952, P. aequalis Haupt, 1952; P. auriguttata Burmeister, 1872 = P. aurimacula Mocsáry, 1885, P. flavicornis Mocsáry, 1894, P. guttata Lucas, 1895, P. incendiaria Lucas, 1895, P. pubiventris Lucas, 1895, P. planifrons Lucas, 1895, P. lestes Lucas, 1895, P. villosa Brèthes, 1908; P. sabina Mocsáry, 1885 = P. astioles Banks, 1946; and P. purpureipes Packard, 1869 = P. chlorana Mocsáry, 1885, P. antennalis Cameron, 1893, P. sulcifrons Cameron, 1903, P. carinata Brèthes, 1914, P. equatoriana Brèthes, 1914, P. angusta Banks, 1946. Keys to all forms are given. The mimicry-groups of P. atripennis Fabricius, 1804, and P. completa Smith, 1855, are defined and described and a comparative account of mimicry based on all four mimicrygroups in Pepsis is given. Lists of excluded species (with their current taxonomic placement and depository where ascertained), unplaced names, and a nomen nudum are given.