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Program provides extrapolation calculations of high gas temperature based on theoretical heating curve of pulsed thermocouples. Program requires as input mach number, wall temperature, and total pressure in addition to thermocouple data. Tests indicate program extrapolates reasonably-accurate gas temperatures from pulsed-thermocouple data.
A gas temperature measurement system with compensated frequency response of 1 kHz and capability to operate in the exhaust of a gas turbine engine combustor was developed. A review of available technologies which could attain this objective was done. The most promising method was identified as a two wire thermocouple, with a compensation method based on the responses of the two different diameter thermocouples to the fluctuating gas temperature field. In a detailed design of the probe, transient conduction effects were identified as significant. A compensation scheme was derived to include the effects of gas convection and wire conduction. The two wire thermocouple concept was tested in a laboratory burner exhaust to temperatures of about 3000 F and in a gas turbine engine to combustor exhaust temperatures of about 2400 F. Uncompensated and compensated waveforms and compensation spectra are presented.
A temperature measurement system, which allows the simultaneous sampling of up to 80 separate thermocouples, was developed. The minimum resolution for the system corresponds to + or - 0.16 C per least significant bit of the A/D converter. The time constant values lambda, for each of the 64 thermocouples, were determined experimentally at 7 mps. Software routines were used to correct the measured temperatures for the effect of lambda for each thermocouple. The temperature measurement system was utilized to study the thermal field of a heated jet discharging perpendicularly into a low and a high disturbance level cross stream for a given momentum flux ratio and for three overheated values. The peak instantaneous temperatures reveal that strong molecular diffusion was operative. Various measures of the thermal field, for the disturbed case, suggest that the jet column remains relatively compact while being buffeted by the ambient turbulence field and that its penetration, into the cross wind, is inhibited by the presence of the strong disturbance field.
The hot section technology (HOST) dynamic gas temperature measurement system computer program acquires data from two type B thermocouples of different diameters. The analysis method determines the in situ value of an aerodynamic parameter T, containing the heat transfer coefficient from the transfer function of the two thermocouples. This aerodynamic parameter is used to compute a fequency response spectrum and compensate the dynamic portion of the signal of the smaller thermocouple. The calculations for the aerodynamic parameter and the data compensation technique are discussed. Compensated data are presented in either the time or frequency domain, time domain data as dynamic temperature vs time, or frequency domain data.
Thermocouple and resistance thermometer expand thermometer scale. Switch thrown up to connect platinum resistance temperature detector or down to connect (platinum/rhodium)/platinum thermocouple to meter. Thermocouple integral part of platinum resistance temperature detector wiring.
Flame temperatures in nonsooting and sooting environments were successfully measured by radiometry for pre-mixed propane-oxygen laminar flames stabilized on a water-cooled, porous sintered-bronze burner. The measured temperatures in the nonsooting flames were compared with fine-wire thermocouple measurements. The results show excellent agreement below 1700 K, and when the thermocouple measurements were corrected for radiation effects, the agreement was good for even higher temperatures. The benefits of radiometry are: (1) the flow is not disturbed by an intruding probe, (2) calibration is easily done using a blackbody source, and (3) measurements can be made even with soot present. The theory involved in the radiometry measurements and the energy balance calculations used to correct the thermocouple temperature measurements are discussed.
A plug-type heat flux gauge formed in a material specimen and having a thermoplug integrally formed in the material specimen, and a method for making the same are disclosed. The thermoplug is surrounded by a concentric annulus, through which thermocouple wires are routed. The end of each thermocouple wire is welded to the thermoplug, with each thermocouple wire welded at a different location along the length of the thermoplug. The thermoplug and concentric annulus may be formed in the material specimen by electrical discharge machining and trepanning procedures.
A method of making a plug-type heat flux gauge in a material specimen in which a thermoplug is integrally formed in the specimen is disclosed. The thermoplug and concentric annulus are formed in the material specimen by electrical discharge machining and trepanning procedures. The thermoplug is surrounded by a concentric annulus through which thermocouple wires are routed. The end of each thermocouple wire is welded to the thermoplug, with each thermocouple wire welded at a different location along the length of the thermoplug.
In proposed technique for determining time-varying temperature of flowing gas, raw measurements of three thermocouples of different sizes processed by relatively simple data-reduction software. Three-thermocouple technique overcomes limitation of single-thermocouple technique.
Thin-film heat-flux sensors have been constructed in the form of arrays of thermocouples on upper and lower surfaces of an insulating layer, so that flux values are proportional to the temperature difference across the upper and lower surface of the insulation material. The sensor thermocouples are connected in thermopile arrangement, and the structure is patterned with photolithographic techniques. Both chromel-alumel and Pt-Pt/Rh thermocouples have been devised; the later produced 28 microvolts when exposed to the radiation of a 1000 C furnace.
A method that makes use of thermocouples has been developed to measure the temperature of engine pistons operating at high speeds. The thermocouples installed on the moving piston are connected with a potentiometer outside the engine by means of pneumatically operated plungers, which make contact with the piston thermocouples for about 10 crankshaft degrees at the bottom of the piston stroke. The equipment is operated satisfactory at engine speeds of 2,400 r.p.m. and shows promise of successful operation at higher engine speeds. Measurements of piston temperatures in a liquid-cooled compression-ignition engine and in an air-cooled spark-ignition are presented.
The accuracy and durability of two means for measuring the metal temperature of gas-turbine blades were investigated experimentally in a turbojet engine. Cemented thermocouples that could be placed in walls as thin 0.010 inch and commercial temperature-indicating paints were evaluated at turbine-inlet temperatures up to about 1650 degrees F and for centrifugal forces equivalent to about 42,000 times gravity. The cemented thermocouples operated satisfactorily. The temperature-indicating paints were neither as accurate nor as durable as the cemented thermocouples.
A method is described by which the total temperature of the gases at the combustion-chamber outlet of a ram-jet engine may be determined from the loss in total pressure measured across the combustion chamber. A working chart is presented by means of which the ratio of the total temperature of the gases at the combustion-chamber outlet to the total temperature of the gases at the combustion-chamber inlet may be determined from the measured loss of total pressure across the combustion chamber and the known values of air flow, total pressure, and total temperature at the combustion-chamber inlet. Values of total-temperature ratio across the combustion chamber of a 20-inch ram jet were obtained in the Cleveland altitude wind tunnel over a range of pressure altitudes from 6000 to 15,000 feet. The difference between the temperature ratio across the combustion chamber determined from the chart and that obtained from the thermocouple measurement was within 6.2 percent of the thermocouple-temperature ratio and was within the accuracy of the thermocouple measurements.
A temperature sensor uses a type R thermocouple wire element in a ceramic sheath to sense temperatures up to 3,200 deg F., and is particularly suitable for flexible insulations. The sensor includes a thermocouple wire embedded in a sheath having two sections disposed at right angles to each other. The junction of the thermocouple is located at one end of one of the sections and the lead wires extend from the other section. The section which includes the junction is secured to a flexible surface with ceramic cement.
Rig for testing ball bearings under radial and axial loads and measuring surface temperatures undergoing development. Includes extensible thermocouples: by means of bellows as longitudinal positioners, thermocouples driven into contact with bearing balls to sense temperatures immediately after test run. Not necessary to disassemble rig or to section balls to obtain indirect indications of maximum temperatures reached. Thermocouple measurements indicate temperatures better than temperature-sensitive paints.
In order to validate the benefits of high aspect ratio cooling channels in a large scale rocket combustion chamber, a high pressure, 89 kN (20,000 lbf) thrust, contoured combustion chamber was tested in the NASA Lewis Research Center Rocket Engine Test Facility. The combustion chamber was tested at chamber pressures from 5.5 to 11.0 MPa (800-1600 psia). The propellants were gaseous hydrogen and liquid oxygen at a nominal mixture ratio of six, and liquid hydrogen was used as the coolant. The combustion chamber was extensively instrumented with 30 backside skin thermocouples, 9 coolant channel rib thermocouples, and 10 coolant channel pressure taps. A total of 29 thermal cycles, each with one second of steady state combustion, were completed on the chamber. For 25 thermal cycles, the coolant mass flow rate was equal to the fuel mass flow rate. During the remaining four thermal cycles, the coolant mass flow rate was progressively reduced by 5, 6, 11, and 20 percent. Computer analysis agreed with coolant channel rib thermocouples within an average of 9 percent and with coolant channel pressure drops within an average of 20 percent. Hot-gas-side wall temperatures of the chamber showed up to 25 percent reduction, in the throat region, over that of a conventionally cooled combustion chamber. Reducing coolant mass flow yielded a reduction of up to 27 percent of the coolant pressure drop from that of a full flow case, while still maintaining up to a 13 percent reduction in a hot-gas-side wall temperature from that of a conventionally cooled combustion chamber.
The Microgravity Smoldering Combustion (MSC) experiment is a study of the smolder characteristics of porous combustible materials in a microgravity environment. The objective of the study is to provide a better understanding of the controlling mechanisms of smolder, both in microgravity and normal earth gravity. Experiments have been conducted aboard the NASA Space Shuttle in the Get Away Special Canister (GAS-CAN), an apparatus requiring completely remote operation. Future GAS-CAN experiments will utilize an ultrasound imaging system (UIS) which has been incorporated into the MSC experimental apparatus. Thermocouples are currently used to measure temperature and reaction front velocities. A less intrusive method is desirable, however, as smolder is a very weak reaction and it has been found that heat transfer along the thermocouple is sufficient to affect the smolder reaction. It is expected that the UIS system will eventually replace the existing array of thermocouples as a non-intrusive technique without compromising data acquisition. The UIS measures line of sight permeability, providing information about the reaction front position and extent. Additionally, the ignition sequence of the MSC experiments has been optimized from previous experiments to provide longer periods of self-supported smolder. An ignition protocol of a fixed power to the igniter for a fixed time is now implemented. This, rather than a controlled temperature profile ignition protocol at the igniter surface, along with the UIS system, will allow for better study of the effect of gravity on a smolder reaction.
One of the goals of the Engineering Test Motor (ETM-3) static test was to verify analytical models through the use of instrumentation that provide real-time transient temperature response of ablating phenolic liners. Accurate measurement of in-depth temperature is critical for validating the analytical models and assessing design safety margins for nozzle insulation materials. Recent developments of in-depth thermocouple plugs have been made to more accurately measure temperature response of nozzle Liners. Thermocouple plugs were installed at two axial stations (four plugs circumferentially at each station) near the end of the ETM-3 nozzle aft exit cone to gather temperature histories during and after motor operation. The thermocouple plugs were placed at depths that bounded the reusable solid rocket motor nominal measured char depth in order to portray the carbon phenolic temperature response during the charring process. The data were used to verify the analytical models during motor operation and to better define char penetration during heat soak after motor burn out.