Gregory J. Brauckmann

Bio: Gregory J. Brauckmann is an academic researcher from Langley Research Center. The author has contributed to research in topics: Mach number & Hypersonic speed. The author has an hindex of 13, co-authored 22 publications receiving 437 citations.

Shuttle Orbiter Experimental Boundary-Layer Transition Results with Isolated Roughness

Abstract: The effect of isolated roughness on the windward surface boundary layer of the Shuttle Orbiter has been experimentally examined in the NASA Langley Research Center 20-InchMach 6 Tunnel. The size and location of isolated roughness elements (intended to simulate raised ormisalignedShuttleOrbiter Thermal Protection System tiles and protruding gap Ž ller material) were varied to systematically examine the response of the boundary layer. Global heat transfer images of the windward surface of a 0.75%-scaleOrbiter at an angle of attack of 40 deg were obtained over a range of Reynolds numbers using phosphor thermography and were used to infer the status of the boundary layer. Computationalpredictions were performed to provide both laminar and turbulent heating levels for comparison to the experimental data and to provide  owŽ eld parameters used for investigatingboundary-layer transition correlations. A variety of roughness heights and locations along the windward centerline were used. The roughness-transition correlation, using the predicted edge parameters Re /Me and k/ , was well behaved. The off-centerline results illustrate the potential for an asymmetric transition pattern to be isolated to one side of the vehicle, thereby causing the increased yawing moments experienced in  ight.

Aerodynamics of Stardust Sample Return Capsule

Abstract: Successful return of interstellar dust and cometary material by the Stardust Sample Return Capsule requires an accurate description of the Earth entry vehicle''s aerodynamics. This desciption must span the hypersonic-rarefied, hypersonic-continuum, supersonic, transonic, and subsonic flow regimes. Data from numerous sources are compiled to accomplish this objective. These include Direct Simulation Monte Carlo analyses, thermochemical nonequilibrium computational fluid dynamics, transonic computational fluid dynamics, existing wind tunnel data, and new wind tunnel data. Four observations are highlighted: 1) a static instability is revealed in the free-molecular and early transitional-flow regime due to aft location of the vehicle''s center-of-gravity, 2) the aerodynamics across the hypersonic regime are compared with the Newtonian flow approximation and a correlation between the accuracy of the Newtonian flow assumption and the sonic line position is noted, 3) the primary effect of shape change due to ablation is shown to be a reduction in drag, and 4) a subsonic dynamic instability is revealed which will necessitate either a change in the vehicle''s center-of-gravity location or the use of a stabilizing drogue parachute.

Aerodynamic Characteristics, Database Development and Flight Simulation of the X-34 Vehicle

Abstract: An overview of the aerodynamic characteristics, development of the preflight aerodynamic database and flight simulation of the NASA/Orbital X-34 vehicle is presented in this paper. To develop the aerodynamic database, wind tunnel tests from subsonic to hypersonic Mach numbers including ground effect tests at low subsonic speeds were conducted in various facilities at the NASA Langley Research Center. Where wind tunnel test data was not available, engineering level analysis is used to fill the gaps in the database. Using this aerodynamic data, simulations have been performed for typical design reference missions of the X-34 vehicle.

Visualization of a Capsule Entry Vehicle Reaction-Control System Thruster

Abstract: Planar laser-induced fluorescence (PLIF) was used to visualize the reaction control system (RCS) jet flow emanating from the aft-body of an Apollo-geometry capsule test article in the NASA Langley Research Center 31-Inch Mach 10 Air wind tunnel. The RCS jet was oriented normal to the aft surface of the model and had a nominal Mach number of 2.94. The composition of the jet gas by mass was 95% nitrogen (N2) and 5% nitric oxide (NO). The RCS jet flowrate varied between zero and 0.5 standard liters per minute and the angle of attack and tunnel stagnation pressure were also varied. PLIF was used to excite the NO molecules for flow visualization. These flow visualization images were processed to determine the trajectory and to quantify the flapping of the RCS jet. The spatial resolution of the jet trajectory measurement was about 1 mm and the single-shot precision of the measurement was estimated to be 0.02 mm in the far field of the jet plume. The jet flapping, measured by the standard deviation of the jet centerline position was as large as 0.9 mm, while the jet was 1.5-4 mm in diameter (full width at half maximum). Schlieren flow visualization images were obtained for comparison with the PLIF. Surface pressures were also measured and presented. Virtual Diagnostics Interface (VIDI) technology developed at NASA Langley was used to superimpose and visualize the data sets. The measurements demonstrate some of the capabilities of the PLIF method while providing a test case for computational fluid dynamics (CFD) validation.

Experimental and Computational Analysis of Shuttle Orbiter Hypersonic Trim Anomaly

Abstract: During the high-Mach-number, high-altitude portion of the first entry of the Shuttle Orbiter, the vehicle exhibited a nose-up pitching moment relative to preflight prediction of approximately Delta Cm = 0.03. This trim anomaly has been postulated to be due to compressibility, viscous, and/or real-gas (lowered specific heat ratio gamma) effects on basic body pitching moment, body-flap effectiveness, or both. In order to assess the relative contribution of each of these effects, an experimental study was undertaken to examine the effects of Mach number, Reynolds number, and ratio of specific heats. Complementary computational solutions were obtained for wind-tunnel and flight conditions. The primary cause of the anomaly was determined to be lower pressures on the aft windward surface of the Orbiter than deduced from hypersonic wind-tunnel tests with ideal- or near-ideal-gas test flow. The lower pressure levels are a result of the lowering of the flowfield gamma due to high-temperature effects. This phenomenon was accurately simulated in a hypersonic wind tunnel using a heavy gas, which provided a lower, gamma, and was correctly predicted by Navier-Stokes computations using nonequilibrium chemistry.

Hypersonic vehicle simulation model: Winged-cone configuration

Abstract: Aerodynamic, propulsion, and mass models for a generic, horizontal-takeoff, single-stage-to-orbit (SSTO) configuration are presented which are suitable for use in point mass as well as batch and real-time six degree-of-freedom simulations The simulations can be used to investigate ascent performance issues and to allow research, refinement, and evaluation of integrated guidance/flight/propulsion/thermal control systems, design concepts, and methodologies for SSTO missions Aerodynamic force and moment coefficients are given as functions of angle of attack, Mach number, and control surface deflections The model data were estimated by using a subsonic/supersonic panel code and a hypersonic local surface inclination code Thrust coefficient and engine specific impulse were estimated using a two-dimensional forebody, inlet, nozzle code and a one-dimensional combustor code and are given as functions of Mach number, dynamic pressure, and fuel equivalence ratio Rigid-body mass moments of inertia and center of gravity location are functions of vehicle weight which is in turn a function of fuel flow

Aerothermodynamics of the Stardust Sample Return Capsule

Abstract: The development of a new high-Ž delity methodology for predicting entry  ows with coupled radiation and ablation is described. The prediction methodology consists of an axisymmetric, nonequilibrium, Navier–Stokes  ow solver loosely coupled to a radiation prediction code and a material thermal response code. The methodology is used to simulate the 12.6-km/s Earth atmospheric entry of the Stardust sample return capsule using ablating and nonablating boundary conditions. These  ow simulations are used to size and design the Stardust forebody and afterbody heatshields and develop arcjet test conditions and models. The  ow simulations indicate that the afterbody heating and pressure proŽ les in time are signiŽ cantly different than the forebody heating and pressure proŽ les. This result is explained in terms of the pertinent aerothermodynamicsof the  owŽ eld and the vehicle’s geometry.When applied to the afterbody thermal protection system, these results show that the traditional afterbody heatshield design approach is nonconservative for the Stardust sample return capsule shape and entry conditions.

Multi-Component Diffusion With Application to Computational Aerothermodynamics

Abstract: The accuracy and complexity of solving multi-component gaseous diffusion using the detailed multi-component equations, the Stefan-Maxwell equations, and two commonly used approximate equations have been examined in a two part study. Part I examined the equations in a basic study with specified inputs in which the results are applicable for many applications. Part II addressed the application of the equations in the Langley Aerothermodynamic Upwind Relaxiation Algorithm (LAURA) computational code for high-speed entries in Earth''s atmosphere. The results showed that the presented iterative scheme for solving the Stefan-Maxwell equations is an accurate and effective method as compared with solutions of the detailed equations. In general, good accuracy with the approximate equations cannot be guaranteed for a species or all species in a multi-component mixture. "Corrected" forms of the approximate equations that ensured the diffusion mass fluxes sum to zero, as required, were more accurate than the uncorrected forms. Good accuracy, as compared with the Stefan-Maxwell results, were obtained with the "corrected" approximate equations in defining the heating rates for the three Earth entries considered in Part II.

Review and synthesis of roughness-dominated transition correlations for reentry applications

Abstract: Nomenclature a = constant; Fig. 1 C , C 0 = constants k = roughness element height, ft N k = average roughness element height, ft L = vehicle length, ft M = Mach number n = exponent; Fig. 1 N R = Poll’s transition parameter; Eq. (1) Reke = roughnessReynolds number based on height k and edge conditions Rekk = roughnessReynolds number based on height k and conditions at k Reμ = Reynolds number based on height μ and edge conditions U = velocity component parallel to test surface or velocity component perpendicularto attachment line, ft /s V = velocity component parallel to attachment line, ft/s X = generalized disturbanceparameter or axial coordinate along windward centerline x = coordinate perpendicular to attachment line, ft Y = generalized transition parameter ® = angle of attack, deg ± = smooth-wall laminar boundary-layer thickness, ft = Poll’s length scale [Eq. (2)], ft μ = smooth-wall laminar boundary-layermomentum thickness, ft 1 = viscosity, lbm/ft ¢ s o = kinematic viscosity, ft2/s 1⁄2 = density, lbm/ft

Hypersonic Boundary-Layer Trip Development for Hyper-X

Abstract: Boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels.Fivedifferenttripconegurationswerecomparedinthreehypersonicfacilities:theNASALangleyResearch Center 20-Inch Mach 6 Air and 31-Inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory. Heat-transfer distributions, utilizing the phosphor thermography and thin-elm techniques, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles of attack of 0, 2, and 4 deg; Reynolds numbers based on model length of 1.2 ££ 10 6‐15.4 £ 10 6 ; and inlet cowl door simulated in both open and closed positions. Comparisons of boundary-layer transition as a result of discrete roughness elements have led to the selection of a trip coneguration for the Hyper-X Mach 7 eight vehicle.