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

Reduction of data from thin-film heat-transfer gages - A concise numerical technique.

Determination of heat-transfer rates from transient surface temperature measurements

A performance analysis of gas-driven shock tubes shows that transonic airfoil flows with chord Reynolds numbers in the range of 100 million can be generated behind the primary shock in a large shock tube. A study of flow over simple airfoils has been carried out at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photos and airfoil pressure measurements show that steady transonic flows similar to those observed for the airfoils in wind tunnels can be generated within the available testing time in a shock tube with either properly-contoured test section walls or a properly-designed slotted-wall test section. The study indicates that the shock tube is a useful facility for studying two-dimensional high Reynolds number transonic airfoil flows.

Shock Tube as a Device for Testing Transonic Airfoils at High Reynolds Numbers

Experimentally determined response characteristics of two types or hot-element wall shear stress gages are described for unsteady turbulent flows. One is the substrate-type thin-film gage and the other is a gage with a small cavity below a surface-level thin-film quartz-fiber element. Two substrate-type gages were tested: a quartz-substrate gage, and a glue-on gage with a thin flexible substrate. Steady-flow gage calibrations were used in reducing unsteady flow data. The cavity gage showed excellent response to a steplike flow, whereas the two substrate gages exhibited response times on the order of tens of seconds for the same flow. Steady-flow shear stress turbulence levels indicated by the cavity and substrate gages were 28 and 12%, respectively

Response of hot-element wall shear stress gages in unsteady turbulent flows

An experimental investigation of the time-dependent response of hot element wall shear stress gages in unsteady periodic air flows is reported. The study has focused on wall shear stress in laminar oscillating flows produced on a flat plate by a free stream velocity composed of a mean component and a superposed sinusoidal variation. Two types of hot element gages, platinum film and flush wire, were tested for values of reduced frequency ranging from 0.14 to 2.36. Values of the phase angle of the wall shear stress variation relative to the free stream velocity, as indicated by the hot element gages, are compared with numerical prediction. The comparisons show that the gages indicate a wall shear stress variation that lags the true variation, and that the gages will also not indicate the correct wall shear stress variation in periodic turbulent flows.

W. J. Cook

Papers

Reduction of data from thin-film heat-transfer gages - A concise numerical technique.

Determination of heat-transfer rates from transient surface temperature measurements

Shock Tube as a Device for Testing Transonic Airfoils at High Reynolds Numbers

Response of hot-element wall shear stress gages in unsteady turbulent flows

Response of hot-element wall shear-stress gages in laminar oscillating flows