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.

Advances in Heat Flux Measurements

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.

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Hypersonic Boundary-Layer Trip Development for Hyper-X

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.

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Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

Experimental and numerical investigation of 2-D backward-facing step flow

The objective of this report is to provide useful engineering formulations and to instill a modest degree of physical understanding of the phenomena governing convective aerodynamic heating at high flight speeds. Some physical insight is not only essential to the application of the information presented here, but also to the effective use of computer codes which may be available to the reader. A discussion is given of cold-wall, laminar boundary layer heating. A brief presentation of the complex boundary layer transition phenomenon follows. Next, cold-wall turbulent boundary layer heating is discussed. This topic is followed by a brief coverage of separated flow-region and shock-interaction heating. A review of heat protection methods follows, including the influence of mass addition on laminar and turbulent boundary layers. Also discussed are a discussion of finite-difference computer codes and a comparison of some results from these codes. An extensive list of references is also provided from sources such as the various AIAA journals and NASA reports which are available in the open literature.

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A review of high-speed, convective, heat-transfer computation methods

Further Experiments on Shock Tube Wall Boundary-Layer Transition

The unsteady shock‐induced laminar boundary layer on a flat plate is theoretically studied for shock speeds ranging from 1.12 to 9 km/sec. Boundary layer flows for real air, assumed to be in thermochemical equilibrium, are analyzed by extending Lam's work on shock‐induced laminar boundary layers to equilibrium dissociated and ionized flows. A complete description of the unsteady nature of the boundary layer is presented in terms of heat transfer and several boundary layer thickness quantities as functions of a single time‐position variable. Boundary layer development is considered for two points of view. Time‐dependent boundary layer development and approach to steady state for any fixed position on the plate is described, as is the configuration of the boundary layer with position on the plate at any point in time. Quantities presented in graphical form permit accurate values for descriptive boundary layer quantities to be easily obtained for post shock pressures at or near 1 atm. Results indicate that in shock‐tube experiments where observations of the flow are made through a boundary layer that influences experimental measurements, it is desirable to employ a flat plate to generate a boundary layer that minimizes the boundary layer thickness and diminishes or eliminates boundary layer unsteadiness.

Characteristics of the Unsteady Shock‐Induced Laminar Boundary Layer on a Flat Plate

The shock tube as a device to fulfill current needs for testing of transonic airfoils at high Reynolds numbers is considered. Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers in the range of 100 million can be produced. A study of flow development over a simple airfoil has been carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls. The study indicates that the shock tube is a viable alternative as a facility for studying high Reynolds number transonic airfoil flows.

Use of shock tubes in high Reynolds number transonic testing

Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.

https://ntrs.nasa.gov/api/citations/19790003842/downloads/19790003842.pdf

Application of shock tubes to transonic airfoil testing at high Reynolds numbers

Changes in the time-dependent boundary-layer turbulence in an oscillating flow were investigated using the oscillating-flow facility described by Cook et al. (1985). Hot wire measurements were made across the boundary layer for frequencies ranging from 4 to 21 Hz at various positions from the boundary layer origin and with different amplitudes of velocity oscillation. The results clearly demonstrate that the turbulence level is influenced by the imposed oscillation. The variation of the turbulence level is not predicted well by the variation in boundary-layer thickness alone, indicating that the interaction between the imposed oscillation and the turbulence structure is more complex. It is suggested that turbulence models developed for steady flow may not be adequate for periodic turbulent flows.

W. J. Cook

Papers

Further Experiments on Shock Tube Wall Boundary-Layer Transition

Characteristics of the Unsteady Shock‐Induced Laminar Boundary Layer on a Flat Plate

Use of shock tubes in high Reynolds number transonic testing

Application of shock tubes to transonic airfoil testing at high Reynolds numbers

Measurements of time-dependent boundary layer turbulence in an oscillating flow