The Characterization of a High Heat Flux Gardon Gauge
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McLaughlin, Luke Patrick↗
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Development of Readily Available & Robust High Heat Flux Gardon Gauges
Concentrated solar power (CSP) technologies deliver concentrated solar energy as a heat source to industrial processes, power generation cycles, and chemical cycles. CSP systems require accurate and reliable high flux measurements, and next generation CSP systems will require flux measurement up to 1000 W/cm2. Existing flux measurement devices do not comprehensively meet the flux rating, cycle life, cost, and lead-time needs of stakeholders, necessitating the development of an improved flux sensor. In this study, Sandia National Laboratories (SNL) partnered with Hukseflux Thermal Sensors to develop a low-cost, short lead-time, and robust flux sensor rated to 250 W/cm2. Three prototype circular foil gauge designs were assessed for performance at the National Solar Thermal Test Facility (NSTTF) at SNL. Each gauge design measured flux up to 250 W/cm2 with <5% measurement error. Following baseline error quantification, gauges were exposed to flux above 500 W/cm2 to assess gauge failure mechanisms. Gauges physically survived >500 W/cm2 flux exposure, but measurement error was found to increase after foil coatings reached 400 °C. The results of this study suggest that coating optical properties change at excessive temperatures and that foil coating temperature, rather than heat flux level, dictates the acceptable gauge measurement range.
McLaughlin, Luke (ORCID:0000000303711310)↗
LOFTID Heat Flux Gauge Calibration: What is Truth?
The Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) is a demonstration of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology, which may enable the delivery of heavy payloads to Mars, Venus, and Titan, as well as return to Earth. Unlike rigid aeroshells that are constrained by the size of the rocket’s shroud, inflatable aeroshells can be deployed to a much larger scale, thus allowing a spacecraft to begin its deceleration earlier and experience less heating. On LOFTID, there will be 4 total heat flux gauges (HFG) with a range of 70 W/cm2 and 1 radiometer with a range of 3 W/cm2, arranged as shown in Fig. 1. Both the radiometer and total HFGs are Schmidt-Boelter gauges purchased from an external vendor. Radiative calibrations were performed in-house at NASA Ames’ Sensors and TPS Advanced Research Laboratories (STAR Labs) before and after environmental testing to investigate how the testing affected the sensors' response. Additional rounds of radiative calibration at STAR Labs were also performed in order to investigate the large uncertainties associated with these tests. For example, a survey of multiple calibration facilities concluded that the uncertainty within a given facility was +/-3% [1]. An additional NIST study that calibrated heat flux gauges at 7 different facilities also found the variation in calibration coefficients to be up to ~3% within a given facility, but up to 15% between facilities, suggesting systematic differences between test setups [2]. Finally, the response of heat flux gauges to radiative versus convective heat flux has shown to differ by up to 20% [3],[4]. Because the heat flux gauges on LOFTID will predominantly experience convective heat flux during flight, a convective calibration study was performed at Boeing's Large-Core Arc Tunnel (LCAT) facility. Radiative Calibration Procedure The calibrations performed at STAR Labs utilize a quartz lamp bank (QLB) that provides a maximum heat flux of 50 W/cm2, which bounds the expected LOFTID flight environment. The calibration involves exposing a water-cooled Gardon gauge (reference) and then the unit-under-test (UUT) to 5 different heat fluxes multiple times for 10 seconds each, and then calculating a linear fit. The test setup is shown in Fig. 2. The total HFGs were calibrated at STAR Labs 3 times, denoted as STAR 1 (before environmental testing), STAR 2 (after protoflight vibration and thermal-vacuum testing), and STAR 3 (no change from previous test). All 8 flight-lot total HFGs showed a decrease in full-scale output from STAR 1 to STAR 2 by between 0.5% and 10. The first portion of this investigation was to determine whether the change could be due to differences in temperature between the two calibration runs. A typical linear fit to the calibration data was performed using Eq. 1 where q’ is the heat flux in W/cm2, c is the calibration coefficient, and mV is the sensor output. To account for temperature, the data were fit to a nonlinear function that included both the sensor output (mV) and the temperature from the thermocouple embedded inside the HFG near the surface (T): q'=mV/(c1* T + c0}. The residuals between the fits and the actual data points were calculated for every point, and proven to be much smaller for the temperature-compensated fits than for the linear fits for all sensors. An example is shown in Fig. 3. When the temperature-compensated fits from STAR 1 were applied to the STAR 2 data, the residuals did not improve, suggesting that the change in sensitivity between these two calibration runs was not due to temperature. A third round of calibration (STAR 3) was conducted to further address the temperature dependence of the total HFGs, and the resulting sensitivities matched closely to STAR 2 (within 2%). Temperature-compensated calibration curves were once again fit to the data. In this case, when the temperature-compensated fits from STAR 3 were applied to STAR 2 data, the residuals between the fits and STAR 2 data were much lower than the residuals due to the linear fits. This suggests that the changes seen between STAR 1 and STAR 2 were likely due to actual changes in the sensors caused by the environmental testing between the two calibrations. A modification of the original calibration process, in which the UUT was exposed to each heat flux for just 3 seconds (instead of 10) to reduce the temperature increase during the test, was additionally performed on several of the HFGs. In general, the sensitivities were 1-1.5% lower than from the 10-second tests, but the temperatures were also significantly lower. When the temperature-compensated fits from the 10-second tests were applied to the 3-second test data, the residuals were greatly improved than when just using the linear fits, further suggesting that the temperature-compensated fits may lead to better accuracy than the linear fits in flight. Convective Calibration The second portion of this study was to create a mapping between the radiative and convective calibration coefficients. The majority of the heating during flight will be convective, so it is important to understand how the HFG response differs under these conditions. However, there are no standardized methods for convective calibration [5]. Because the TPS aerothermal response models were validated at LCAT, the same facility was chosen for convective calibration of two of the total HFGs (Fig. 4). Preliminary results showed that the full-scale output was 3% and 8% higher in convective heat flux as compared to radiative heat flux. However, tunnel variation may have contributed to noise and uncertainty in the measurements, and more testing and analysis remains to be done. Scope of Presentation The presentation will include an overview of the changes seen in HFG calibration before and after environmental testing, differences between radiative and convective calibrations, the modeling work done to aid in understanding the sensor response to varying environments, and recommended future work.
H S Alpert↗
Development of advanced high-temperature heat flux sensors
Various configurations of high temperature, heat flux sensors were studied to determine their suitability for use in experimental combustor liners of advanced aircraft gas turbine engines. It was determined that embedded thermocouple sensors, laminated sensors, and Gardon gauge sensors, were the most viable candidates. Sensors of all three types were fabricated, calibrated, and endurance tested. All three types of sensors met the fabricability survivability, and accuracy requirements established for their application.
Atkinson, W. H.↗
Development of advanced high-temperature heat flux sensors. Phase 2: Verification testing
A two-phase program is conducted to develop heat flux sensors capable of making heat flux measurements throughout the hot section of gas turbine engines. In Phase 1, three types of heat flux sensors are selected; embedded thermocouple, laminated, and Gardon gauge sensors. A demonstration of the ability of these sensors to operate in an actual engine environment is reported. A segmented liner of each of two combustors being used in the Broad Specification Fuels Combustor program is instrumented with the three types of heat flux sensors then tested in a high pressure combustor rig. Radiometer probes are also used to measure the radiant heat loads to more fully characterize the combustor environment. Test results show the heat flux sensors to be in good agreement with radiometer probes and the predicted data trends. In general, heat flux sensors have strong potential for use in combustor development programs.
Atkinson, W. H.↗
Turbine blade and vane heat flux sensor development, phase 2
The development of heat flux sensors for gas turbine blades and vanes and the demonstration of heat transfer measurement methods are reported. The performance of the heat flux sensors was evaluated in a cylinder in cross flow experiment and compared with two other heat flux measurement methods, the slug calorimeter and a dynamic method based on fluctuating gas and surface temperature. Two cylinders, each instrumented with an embedded thermocouple sensor, a Gardon gauge, and a slug calorimeter, were fabricated. Each sensor type was calibrated using a quartz lamp bank facility. The instrumented cylinders were then tested in an atmospheric pressure combustor rig at conditions up to gas stream temperatures of 1700K and velocities to Mach 0.74. The test data are compared to other measurements and analytical prediction.
Atkinson, W. H.↗
Development of heat flux sensors for turbine airfoils
The objectives of this program are to develop heat flux sensors suitable for installation in hot section airfoils of advanced aircraft turbine engines and to experimentally verify the operation of these heat flux sensors in a cylinder in a cross flow experiment. Embedded thermocouple and Gardon gauge sensors were developed and fabricated into both blades and vanes. These were then calibrated using a quartz lamp bank heat source and finally subjected to thermal cycle and thermal soak testing. These sensors were also fabricated into cylindrical test pieces and tested in a burner exhaust to verify heat flux measurements produced by these sensors. The results of the cylinder in cross flow tests are given.
Atkinson, William H.↗
Characterization of a 50kW Inductively Coupled Plasma Torch for Testing of Ablative Thermal Protection Materials
With the development of new manned spaceflight capabilities including NASA's Orion capsule and the Space-X Dragon capsule, there is a renewed importance of understanding the dynamics of ablative thermal protection systems. To this end, a new inductively coupled plasma torch facility is being developed at UT-Austin. The torch operates on argon and/or air at plasma powers up to 50 kW. In the present configuration the flow issues from a low-speed subsonic nozzle and the hot plume is characterized using slug calorimetry and emission spectroscopy. Preliminary measurements using emission spectroscopy have indicated that the torch is capable of producing an air plasma with a temperature between 6,000 K and 8,000 K depending on the power and flow settings and an argon plasma with a temperature of approximately 12,000 K. The operation envelope was measured, and heat flux measured for every point within the envelope using both a slug calorimeter and a Gardon gauge heat flux sensor. The torch was found to induce a stagnation point heat flux of between 90 and 225 W/sq cm.
Greene, Benton R.↗
Defining the Operational Envelope for Air Flows in the Miniature Arc-Jet Research Chamber (mARC II)
The second-generation 30 kW miniature Arc-jet Research Chamber (mARC II) at NASA Ames Research Center produces high enthalpy flows relevant for entry systems ground testing. The mARC II facility has recently undergone upgrades, including the installation of a new vacuum system to address the issues preventing it from maintaining underexpanded flow under test conditions. In this work, we present data obtained from an Integrated Systems Testing campaign and provide an initial assessment of arc-jet performance following the upgrades. Air is used as the working gas for the standard mARC II arc-heater configuration with two constrictor disks. Seven runs were investigated for five test conditions to assess the lowest achievable stagnation point heat fluxes for air flow rates of 0.15 or 0.25 g s −1 . The heat flux was measured using a water-cooled Gardon gauge (Ø4.76 mm, 3/16" hemispherical) at 70 mm from the nozzle exit plane. The new vacuum system produced test box pressures in the medium (fine) vacuum range (~0.03 torr, 4 Pa) prior to gas addition and successfully maintained underexpanded flow after gas addition. The upgrade yielded a ~4X reduction in heat flux relative to the previous system for the same set test conditions. We report the lowest heat fluxes measured in mARC II to date, ranging from 26 to 81 W cm −2 , for sonic flow enthalpies of 4–14 MJ kg −1 . Bulk enthalpies estimated using an energy balance method (EB 2 ) are reported for the first time using mARC II.Initial data suggests EB 2 generally estimates lower enthalpies than sonic flow methods for mARC II. Lastly, laminar axisymmetric Navier–Stokes simulations were performed using the NASA DPLR code. Numerical heat flux results show good agreement with experiments at low arc powers (3–15% difference at the minimum set arc current), but discrepancy increases with arc power (49% difference at the maximum set arc current).
Pitot tube↗
Defining the Operational Envelope for Air Flows in the Miniature Arc-Jet Research Chamber (mARC II)
The second-generation 30 kW miniature Arc-jet Research Chamber (mARC II) at NASA Ames Research Center produces high enthalpy flows relevant for entry systems ground testing. The mARC II facility has recently undergone upgrades, including the installation of a new vacuum system to address the issues preventing it from maintaining underexpanded flow under test conditions. In this work, we present data obtained from an Integrated Systems Testing campaign and provide an initial assessment of arc-jet performance following the upgrades. Air is used as the working gas for the standard mARC II arc-heater configuration with two constrictor disks. Seven runs were investigated for five test conditions to assess the lowest achievable stagnation point heat fluxes for air flow rates of 0.15 or 0.25 g s −1 . The heat flux was measured using a water-cooled Gardon gauge (Ø4.76 mm, 3/16" hemispherical) at 70 mm from the nozzle exit plane. The new vacuum system produced test box pressures in the medium (fine) vacuum range (~0.03 torr, 4 Pa) prior to gas addition and successfully maintained underexpanded flow after gas addition. The upgrade yielded a ~4X reduction in heat flux relative to the previous system for the same set test conditions. We report the lowest heat fluxes measured in mARC II to date, ranging from 26 to 81 W cm −2 , for sonic flow enthalpies of 4–14 MJ kg −1 . Bulk enthalpies estimated using an energy balance method (EB 2 ) are reported for the first time using mARC II.Initial data suggests EB 2 generally estimates lower enthalpies than sonic flow methods for mARC II. Lastly, laminar axisymmetric Navier–Stokes simulations were performed using the NASA DPLR code. Numerical heat flux results show good agreement with experiments at low arc powers (3–15% difference at the minimum set arc current), but discrepancy increases with arc power (49% difference at the maximum set arc current).
arc-jet↗
Expanding the Measurement Capabilities of the mARC II Arc-Jet to Map the Operating Envelope for High-Enthalpy Air Flows
The mARC II is a 30 kW arc-jet facility at NASA Ames Research Center developed to produce high-enthalpy flows for low-cost technology development purposes. In this work, we introduce new measurement capabilities following the latest facility upgrades and begin characterizing the operating envelope for the two-disk arc-heater configuration with air as the working gas. The lower bound of the envelope corresponds to the lowest set current and the farthest sensor distance from the nozzle (𝐼set = 40 A, 𝑧 = 69 mm), while the upper bound corresponds to the highest current and closest sensor distance (𝐼set = 200 A, 𝑧 = 2 mm). Stagnation point heat flux was measured using a water-cooled Gardon gauge (⌀4.76 mm, 3/16" hemispherical), with values ranging from 15 ≤ q̇₀ ≤ 900 W/cm2. Stagnation pressure was measured using a water-cooled Pitot probe (⌀4.76 mm, 3/16" hemispherical), with values ranging from 18 ≤ p₀ ≤ 1100 Pa. The upgraded vacuum system is demonstrated to significantly extend the lower end of the operating envelope of the mARC II. Additionally, measurements of bulk enthalpy (3 ≤ h ≤ 13 MJ/kg) and stagnation enthalpy (4 ≤ h₀ ≤ 34 MJ/kg) demonstrate the mARC II facility's high-enthalpy capabilities. Completion of the operating envelope characterization is ongoing in tandem with the development of complementary diagnostic capabilities and numerical simulations.
TPS↗
Expanding the Measurement Capabilities of the mARC II Arc-Jet to Map the Operating Envelope for High-Enthalpy Air Flows
The mARC II is a 30 kW arc-jet facility at NASA Ames Research Center developed to produce high-enthalpy flows for low-cost technology development purposes. In this work, we introduce new measurement capabilities following the latest facility upgrades and begin characterizing the operating envelope for the two-disk arc-heater configuration with air as the working gas. The lower bound of the envelope corresponds to the lowest set current and the farthest sensor distance from the nozzle (𝐼set = 40 A, 𝑧 = 69 mm), while the upper bound corresponds to the highest current and closest sensor distance (𝐼set = 200 A, 𝑧 = 2 mm). Stagnation point heat flux was measured using a water-cooled Gardon gauge (⌀4.76 mm, 3/16" hemispherical), with values ranging from 15 ≤ q̇₀ ≤ 900 W/cm 2 . Stagnation pressure was measured using a water-cooled Pitot probe (⌀4.76 mm, 3/16" hemispherical), with values ranging from 18 ≤ p₀ ≤ 1100 Pa. The upgraded vacuum system is demonstrated to significantly extend the lower end of the operating envelope of the mARC II. Additionally, measurements of bulk enthalpy (3 ≤ h ≤ 13 MJ/kg) and stagnation enthalpy (4 ≤ h₀ ≤ 34 MJ/kg) demonstrate the mARC II facility's high-enthalpy capabilities. Completion of the operating envelope characterization is ongoing in tandem with the development of complementary diagnostic capabilities and numerical simulations.
instrumentation↗
A proposed high-intensity radiometer calibration method using concentrated solar radiation
Accurate calibration of irradiance measurement devices, or radiometers, is essential for ensuring the reliability of measurements in high heat applications such as concentrating solar power (CSP), aerospace, defense, and pulsed power systems. Despite the critical need, existing calibration standards and service providers are limited to irradiance levels below 100 kW/m 2 and specific radiation sources, which is insufficient for many applications. For instance, CSP technologies, particularly those under the Department of Energy’s Solar Energy Technologies Office (SETO) Gen 3 program, require accurate measurements of broadband irradiance at levels exceeding 2000 kW/m 2 . In even more extreme scenarios, such as re-entry vehicles, heat levels can surpass 10000 kW/m 2 . Current ISO standards, specifically ISO 14934–2 and ISO 14934–3, are constrained to lower irradiance levels and dependent on black body heat sources, limiting their applicability for high-intensity broadband irradiance measurements, particularly in concentrated solar applications. Here, to address this shortfall, the National Solar Thermal Test Facility (NSTTF) at Sandia National Laboratories (SNL) proposes a calibration method and facility capable of characterizing radiometers up to 2750 kW/m 2 using concentrated solar irradiance. Calibrating with concentrated sunlight is important for solar applications as it aligns the calibration process with the solar spectrum. This alignment is crucial for minimizing systematic errors and avoiding the need for additional corrections that may arise when radiometers designed for solar applications are calibrated using black-body or electrical sources. This paper presents the present day NSTTF characterization facility and procedure, detailing the proposed calibration method and uncertainty quantification. The presented method builds upon 1980′s NSTTF methodology and involves both theoretical and empirical methods to establish a robust relationship between gauge voltage output and irradiance intensity, quantifying both measurement and fitting errors. By addressing the limitations of existing standards and extending the characterization range, this work provides an advancement in the field of high-intensity irradiance measurement and instrumentation characterization.
Gardon gauge↗
Local Heat Flux Measurements with Single Element Coaxial Injectors
To support the mission for the NASA Vision for Space Exploration, the NASA Marshall Space Flight Center conducted a program in 2005 to improve the capability to predict local thermal compatibility and heat transfer in liquid propellant rocket engine combustion devices. The ultimate objective was to predict and hence reduce the local peak heat flux due to injector design, resulting in a significant improvement in overall engine reliability and durability. Such analyses are applicable to combustion devices in booster, upper stage, and in-space engines, as well as for small thrusters with few elements in the injector. In this program, single element and three-element injectors were hot-fire tested with liquid oxygen and ambient temperature gaseous hydrogen propellants at The Pennsylvania State University Cryogenic Combustor Laboratory from May to August 2005. Local heat fluxes were measured in a 1-inch internal diameter heat sink combustion chamber using Medtherm coaxial thermocouples and Gardon heat flux gauges. Injectors were tested with shear coaxial and swirl coaxial elements, including recessed, flush and scarfed oxidizer post configurations, and concentric and non-concentric fuel annuli. This paper includes general descriptions of the experimental hardware, instrumentation, and results of the hot-fire testing for three of the single element injectors - recessed-post shear coaxial with concentric fuel, flush-post swirl coaxial with concentric fuel, and scarfed-post swirl coaxial with concentric fuel. Detailed geometry and test results will be published elsewhere to provide well-defined data sets for injector development and model validatation.
Jones, Gregg↗
Local Heat Flux Measurements with Single and Small Multi-element Coaxial Element-Injectors
To support NASA's Vision for Space Exploration mission, the NASA Marshall Space Flight Center conducted a program in 2005 to improve the capability to predict local thermal compatibility and heat transfer in liquid propellant rocket engine combustion devices. The ultimate objective was to predict and hence reduce the local peak heat flux due to injector design, resulting in a significant improvement in overall engine reliability and durability. Such analyses are applicable to combustion devices in booster, upper stage, and in-space engines with regeneratively cooled chamber walls, as well as in small thrust chambers with few elements in the injector. In this program, single and three-element injectors were hot-fire tested with liquid oxygen and gaseous hydrogen propellants at The Pennsylvania State University Cryogenic Combustor Laboratory from May to August 2005. Local heat fluxes were measured in a 1-inch internal diameter heat sink combustion chamber using Medtherm coaxial thermocouples and Gardon heat flux gauges, Injector configurations were tested with both shear coaxial elements and swirl coaxial elements. Both a straight and a scarfed single element swirl injector were tested. This paper includes general descriptions of the experimental hardware, instrumentation, and results of the hot-fire testing for three coaxial shear and swirl elements. Detailed geometry and test results the for shear coax elements has already been published. Detailed test result for the remaining 6 swirl coax element for the will be published in a future JANNAF presentation to provide well-defined data sets for development and model validation.
Jones, Gregg↗
Benchmark Wall Heat Flux Data for a GO2/GH2 Single Element Combustor
Wall heat flux measurements in a 1.5 in. diameter circular cross-section rocket chamber for a uni-element shear coaxial injector element operating on gaseous oxygen (GOz)/gaseous hydrogen (GH,) propellants are presented. The wall heat flux measurements were made using arrays of Gardon type heat flux gauges and coaxial thermocouple instrumentation. Wall heat flux measurements were made for two cases. For the first case, GOZ/GHz oxidizer-rich (O/F=l65) and fuel-rich preburners (O/F=1.09) integrated with the main chamber were utilized to provide vitiated hot fuel and oxidizer to the study shear coaxial injector element. For the second case, the preburners were removed and ambient temperature gaseous oxygen/gaseous hydrogen propellants were supplied to the study injector. Experiments were conducted at four chamber pressures of 750, 600, 450 and 300psia for each case. The overall mixture ratio for the preburner case was 6.6, whereas for the ambient propellant case, the mixture ratio was 6.0. Total propellant flow was nominally 0.27-0.29 Ibm/s for the 750 psia case with flowrates scaled down linearly for lower chamber pressures. The axial heat flux profile results for both the preburner and ambient propellant cases show peak heat flux levels a t axial locations between 2.0 and 3.0 in. from the injector face. The maximum heat flux level was about two times greater for the preburner case. This is attributed to the higher injector fuel-to-oxidizer momentum flux ratio that promotes mixing and higher initial propellant temperature for the preburner case which results in a shorter reaction zone. The axial heat flux profiles were also scaled with respect to the chamber pressure to the power 0.8. The results at the four chamber pressures for both cases collapsed to a single profile indicating that at least to first approximation, the basic fluid dynamic structures in the flow field are pressure independent as long as the chamber/njector/nozzle geometry and injection velocities remain the same.
Marshall, William M.↗
Numerical study of the effects of boundary conditions on the measurement and calibration of gardon type heat flux sensors
To monitor the high-intensity heat flux conditions that occur in the space shuttle main engine (SSME), it is necessary to use specifically designed heat flux sensors. These sensors, which are of the Gardon-type, are exposed on the measuring face to high-intensity radiative and convective heat fluxes and on the other face to convective cooling. To improve the calibration and measurement accuracy of these gauges, researchers are studing the effect that the thermal boundary conditions have on gauge performance. In particular, they are studying how convective cooling effects the field inside the sensor and the measured heat flux. The first phase of this study involves a numerical study of these effects. Subsequent phases will involve experimental verification. A computer model of the heat transfer around a Garden-type heat flux sensor was developed. Two specific geometries are being considered are: (1) heat flux sensor mounted on a flat-plate; and (2) heat flux sensor mounted at the stagnation point of a circular cylinder. Both of these configurations are representative of the use of heat flux sensors in the components of the SSME. The purpose of the analysis is to obtain a temperature distribution as a function of the boundary conditions.
Krane, M.↗