John R. Micol

Bio: John R. Micol is an academic researcher from Langley Research Center. The author has contributed to research in topics: Mach number & Reynolds number. The author has an hindex of 10, co-authored 17 publications receiving 332 citations.

Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

Abstract: Description, capabilities, recent upgrades, and utilization of the NASA Langley Research Center (LaRC) Aerothermodynamic Facilities Complex (AFC) are presented. The AFC consists of five hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 20, unit Reynolds number from 0.04 to 22 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of three different test gases (air, helium, and tetrafluoromethane), thereby making several of the facilities unique. The Complex represents nearly three-fourths of the conventional (as opposed to impulse)-type hypersonic wind tunnels operational in this country. AFC facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data fr generating the flight aerodynamic databook and final design of the thermal protection system (TPS) (e.g., establishment of flight limitations not to exceed TPS design limits). Modifications and enhancements of AFC hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. Examples illustrating facility utilization in recent years to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs are presented. These programs include Phase I of the Reusable Launch Vehicle (RLV) Advanced Technology Demonstrator, X-33 program, PHase II of the X-33 program, X-34 program, the Hyper-X program ( a Mach 5,7, and 10 airbreathing propulsion flight experiment), and the X-38 program (Experimental Crew Return Vehicle, X-CRV). Current upgrades/enchancements and future plans for the AFC are discussed.

Aerothermodynamic measurement and prediction for a modified orbiter at Mach 6 and 10 in air

Abstract: Detailed heat-transfer rate distributions measured laterally over the windward surface of an orbiter-like configuration using thin-film resistance heat-transfer gauges and globally using the newly developed relative intensity, two-color thermographic phosphor technique are presented for Mach 6 and 10 in air. The angle of attack was varied from 0 to 40 deg, and the freestream Reynolds number based on the model length was varied from 4 x 10(exp 5) to 6 x 10(exp 6) at Mach 6, corresponding to laminar, transitional, and turbulent boundary layers; the Reynolds number at Mach 10 was 4 x 10(exp 5), corresponding to laminar flow. The primary objective of the present study was to provide detailed benchmark heat-transfer data for the calibration of computational fluid-dynamics codes. Predictions from a Navier-Stokes solver referred to as the Langley aerothermodynamic upwind relaxation algorithm and an approximate boundary-layer solving method known as the axisymmetric analog three-dimensional boundary layer code are compared with measurement. In general, predicted laminar heat-transfer rates are in good agreement with measurements.

Aerothermodynamic Measurement and Prediction for Modified Orbiter at Mach 6 and 10

Abstract: Detailed heat-transfer rate distributions measured laterally over the windward surface of an orbiter-like configuration using thin-film resistance heat-transfer gauges and globally using the newly developed relative intensity, two-color thermographic phosphor technique are presented for Mach 6 and 10 in air. The angle of attack was varied from 0 to 40 deg, and the freestream Reynolds number based on the model length was varied from 4 x 10(exp 5) to 6 x 10(exp 6) at Mach 6, corresponding to laminar, transitional, and turbulent boundary layers; the Reynolds number at Mach 10 was 4 x 10(exp 5), corresponding to laminar flow. The primary objective of the present study was to provide detailed benchmark heat-transfer data for the calibration of computational fluid-dynamics codes. Predictions from a Navier-Stokes solver referred to as the Langley aerothermodynamic upwind relaxation algorithm and an approximate boundary-layer solving method known as the axisymmetric analog three-dimensional boundary layer code are compared with measurement. In general, predicted laminar heat-transfer rates are in good agreement with measurements.

Simulation of real-gas effects on pressure distributions for a proposed Aeroassist Flight Experiment vehicle and comparison to prediction

Abstract: Pressure distributions measured on a 60-deg elliptic cone, raked off at a 73-deg angle and having an ellipsoid nose (ellipticity equal to 2.0), are presented for a range of angle of attack from -10 to 15 deg. The high normal shock density ratio aspect of a real gas was simulated by testing in Mach-6 air (normal shock density ratio equal to 5.25) and Mach 6 CF4 (normal shock density ratio equal to 12.0). The effects of Reynolds number, angle of attack, and normal shock density ratio on these measurements are examined, and comparisons are made to an inviscid flowfield computer code known as HALIS. A significant effect of density ratio on pressure distributions on the cone section of the configuration was observed; the magnitude of this effect decreased with increased angle of attack. The effect of Reynolds number on pressures was negligible for forebody pressure distributions, but a measurable effect was noted on base pressures. In general, the HALIS code accurately predicted the measured pressure distributions in air and CF4.

Experimental and predicted aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration at Mach 6 and 10

Abstract: The effects of Mach number, Reynolds number, and ratio of specific heats on the aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration were examined for a range of angles of attack from -5 deg to 50 deg. Predictions made with a Langley-developed, three-dimensional Navier Stokes solver known as LAURA, which was exercised as an Euler solver for the present study, are compared with the experimental results. Unlike the Shuttle Orbiter, which experienced a significant nose-up increment in pitching moment with decreasing specific heat ratio (i.e., real gas effects), the aerodynamic characteristics of this lifting-body configuration are insensitive to changes in specific heat ratio. The maximum trimmed lift-to-drag ratio achieved was about 1.5. Predicted inviscid values of aerodynamic coefficients were generally in good agreement with measurement.

An upwind-biased, point-implicit relaxation algorithm for viscous, compressible perfect-gas flows

Abstract: An upwind-biased, point-implicit relaxation algorithm for obtaining the numerical solution to the governing equations for three-dimensional, viscous, compressible, perfect-gas flows is described. The algorithm is derived using a finite-volume formulation in which the inviscid components of flux across cell walls are described with Roe's averaging and Harten's entropy fix with second-order corrections based on Yee's Symmetric Total Variation Diminishing scheme. Viscous terms are discretized using central differences. The relaxation strategy is well suited for computers employing either vector or parallel architectures. It is also well suited to the numerical solution of the governing equations on unstructured grids. Because of the point-implicit relaxation strategy, the algorithm remains stable at large Courant numbers without the necessity of solving large, block tri-diagonal systems. Convergence rates and grid refinement studies are conducted for Mach 5 flow through an inlet with a 10 deg compression ramp and Mach 14 flow over a 15 deg ramp. Predictions for pressure distributions, surface heating, and aerodynamics coefficients compare well with experiment data for Mach 10 flow over a blunt body.

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

Advances in Heat Flux Measurements

Abstract: Publisher Summary The decade of the 1950s saw great advances in heat-transfer measurement techniques. Optical methods became popular, along with several new heat-flux gages that are still in wide use today, as evidenced by the commercial heat-flux-gage manufacturers. In the last 10 to 20 years, a number of new techniques have been developed and applied, which have greatly increased the resolution and operating range of heat-flux instrumentation. Although the older methods are discussed as the background, the focus of this chapter is on the newer techniques and novel uses of the older techniques. In addition, an effort is made to bring together information from a variety of fields that deal with heat transfer, but that often isn't communicated. Three areas of new capability with important applications are time-resolved heat-flux measurements, simultaneous measurement of spatially distributed heat flux, and heat-flux measurement at high-temperature conditions. Examples of recent advances in these areas are discussed. Such new capabilities, when applied to real-world problems, make the field of heat transfer exciting. From these reviews, it is clear that no one gage or method is good for every application. The limitations of previous heat-flux-gage performance are highlighted and the need for better gage characteristics is expressed. Recent advances have been able to overcome some of these limitations and provide measurements with improved accuracy under conditions previously not possible. The discussion of measurement methods is organized by the first three categories of heat-transfer measurement categories, given in the chapter. Recent measurements using the methods discussed are showcased to demonstrate the advances made in measurement capability. Calibration is briefly addressed in the chapter.

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.

Global Aeroheating Wind-Tunnel Measurements Using Improved Two-Color Phosphor Thermography Method

Abstract: Detailed aeroheating information is critical to the successful design of a thermal protection system (TPS) for an aerospace vehicle. NASA Langley Research Center’ s (LaRC) phosphor thermography method is described. Development of theory is provided for a new weighted two-color relative-intensity e uorescence theory for quantitatively determining surface temperatures on hypersonic wind-tunnel models and an improved application of the one-dimensional conduction theory for use in determining global heating mappings. The phosphor methodology at LaRC is presented including descriptions of phosphor model fabrication, test facilities, and phosphor video acquisition systems. A discussion of the calibration procedures, data reduction, and data analysis is given. Estimates of the total uncertainties (with a 95% cone dence level ) associated with the phosphor technique are shown to be approximately 7 ‐10% in LaRC’ s 31-Inch Mach 10 Tunnel and 8 ‐10% in the 20-Inch Mach 6 Tunnel. A comparison with thin-e lm measurements using 5.08-cm-radius hemispheres shows the phosphor data to be within 7% of thin-e lm measurements and to agree even better with predictions via a LATCH computational e uid dynamics (CFD) solution. Good agreement between phosphor data and LAURA CFD computations on the forebody of a vertical takeoff/vertical lander cone guration at four angles of attack is also shown. In addition, a comparison is givenbetween Mach 6phosphordata andlaminarandturbulentsolutionsgeneratedusing theLAURA,GASP, and LATCH CFD codes on the X-34 cone guration. The phosphor process outlined is believed to provide the aerothermodynamic community with a valuable capability for rapidly obtaining (three to four weeks ) detailed heating information needed in TPS design. Nomenclature A = area of camera array element, m 2 a = effective aperture factor of camera optics, sr b = vehicle wing span from wing tip to wing tip, m C = heat transfer coefe cient constant, h.iw=Tw/ c = specie c heat of model substrate, J/ (kg-K) D = driver constant, iaw.Tw=iw/iTinit F = e ux of light, W/m 2 h = heat transfer coefe cient, kg/ (m 2 -s) I = radiant intensity, W/ (m 2 -sr)