Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “THERMOCOUPLES”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,621 records · Page 90
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Aerospace Sensor Systems: From Sensor Development To Vehicle Application

This paper presents an overview of years of sensor system development and application for aerospace systems. The emphasis of this work is on developing advanced capabilities for measurement and control of aeropropulsion and crew vehicle systems as well as monitoring the safety of those systems. Specific areas of work include chemical species sensors, thin film thermocouples and strain gages, heat flux gages, fuel gages, SiC based electronic devices and sensors, space qualified electronics, and MicroElectroMechanical Systems (MEMS) as well as integrated and multifunctional sensor systems. Each sensor type has its own technical challenges related to integration and reliability in a given application. The general approach has been to develop base sensor technology using microfabrication techniques, integrate sensors with "smart" hardware and software, and demonstrate those systems in a range of aerospace applications. Descriptions of the sensor elements, their integration into sensors systems, and examples of sensor system applications will be discussed. Finally, suggestions related to the future of sensor technology will be given. It is concluded that smart micro/nano sensor technology can revolutionize aerospace applications, but significant challenges exist in maturing the technology and demonstrating its value in real-life applications.

Hunter, Gary W.↗
NASA NTRS2013

RF Reference Switch for Spaceflight Radiometer Calibration

The goal of this technology is to provide improved calibration and measurement sensitivity to the Soil Moisture Active Passive Mission (SMAP) radiometer. While RF switches have been used in the past to calibrate microwave radiometers, the switch used on SMAP employs several techniques uniquely tailored to the instrument requirements and passive remote-sensing in general to improve radiometer performance. Measurement error and sensitivity are improved by employing techniques to reduce thermal gradients within the device, reduce insertion loss during antenna observations, increase insertion loss temporal stability, and increase rejection of radar and RFI (radio-frequency interference) signals during calibration. The two legs of the single-pole double-throw reference switch employ three PIN diodes per leg in a parallel-shunt configuration to minimize insertion loss and increase stability while exceeding rejection requirements at 1,413 MHz. The high-speed packaged diodes are selected to minimize junction capacitance and resistance while ensuring the parallel devices have very similar I-V curves. Switch rejection is improved by adding high-impedance quarter-wave tapers before and after the diodes, along with replacing the ground via of one diode per leg with an open circuit stub. Errors due to thermal gradients in the switch are reduced by embedding the 50-ohm reference load within the switch, along with using a 0.25-in. (approximately equal to 0.6-cm) aluminum prebacked substrate. Previous spaceflight microwave radiometers did not embed the reference load and thermocouple directly within the calibration switch. In doing so, the SMAP switch reduces error caused by thermal gradients between the load and switch. Thermal issues are further reduced by moving the custom, highspeed regulated driver circuit to a physically separate PWB (printed wiring board). Regarding RF performance, previous spaceflight reference switches have not employed high-impedance tapers to improve rejection. The use of open-circuit stubs instead of a via to provide an improved RF short is unique to this design. The stubs are easily tunable to provide high rejection at specific frequencies while maintaining very low insertion loss in-band.

Knuble, Joseph↗
NASA NTRS2013

Quantitative Infrared Image Analysis Of Simultaneous Upstream and Downstream Microgravity Flame Spread over Thermally-Thin Cellulose in Low Speed Forced Flow

The effect of low velocity forced flow on microgravity flame spread is examined using quantitative analysis of infrared video imaging. The objective of the quantitative analysis is to provide insight into the mechanisms of flame spread in microgravity where the flame is able to spread from a central location on the fuel surface, rather than from an edge. Surface view calibrated infrared images of ignition and flame spread over a thin cellulose fuel were obtained along with a color video of the surface view and color images of the edge view using 35 mm color film at 2 Hz. The cellulose fuel samples were mounted in the center of a 12 cm wide by 16 cm tall flow duct and were ignited in microgravity using a straight hot wire across the center of the 7.5 cm wide by 14 cm long samples. Four cases, at 1 atm. 35%O2 in N2, at forced flows from 2 cm/s to 20 cm/s are presented here. This flow range captures flame spread from strictly upstream spread at low flows, to predominantly downstream spread at high flow. Surface temperature profiles are evaluated as a function of time, and temperature gradients for upstream and downstream flame spread are measured. Flame spread rates from IR image data are compared to visible image spread rate data. IR blackbody temperatures are compared to surface thermocouple readings to evaluate the effective emissivity of the pyrolyzing surface. Preheat lengths and pyrolysis lengths are evaluated both upstream and downstream of the central ignition point. A surface energy balance estimates the net heat flux from the flame to the fuel surface along the length of the fuel. Surface radiative loss and gas-phase radiation from soot are measured relative to the net heat feedback from the flame. At high surface heat loss relative to heat feedback, the downstream flame spread does not occur.

Olson, S. L.↗
NASA NTRS2013

Assessment of tbe Performance of Ablative Insulators Under Realistic Solid Rocket Motor Operating Conditions (a Doctoral Dissertation)

Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

Martin, Heath Thomas↗
NASA NTRS2013

Mars Science Laboratory Heatshield Aerothermodynamics: Design and Reconstruction

The Mars Science Laboratory heatshield was designed to withstand a fully turbulent heat pulse based on test results and computational analysis on a pre-flight design trajectory. Instrumentation on the flight heatshield measured in-depth temperatures in the thermal protection system. The data indicate that boundary layer transition occurred at 5 of 7 thermocouple locations prior to peak heating. Data oscillations at 3 pressure measurement locations may also indicate transition. This paper presents the heatshield temperature and pressure data, possible explanations for the timing of boundary layer transition, and a qualitative comparison of reconstructed and computational heating on the as-flown trajectory. Boundary layer Reynolds numbers that are typically used to predict transition are compared to observed transition at various heatshield locations. A uniform smooth-wall transition Reynolds number does not explain the timing of boundary layer transition observed during flight. A roughness-based Reynolds number supports the possibility of transition due to discrete or distributed roughness elements on the heatshield. However, the distributed roughness height would have needed to be larger than the pre-flight assumption. The instrumentation confirmed the predicted location of maximum turbulent heat flux near the leeside shoulder. The reconstructed heat flux at that location is bounded by smooth-wall turbulent calculations on the reconstructed trajectory, indicating that augmentation due to surface roughness probably did not occur. Turbulent heating on the downstream side of the heatshield nose exceeded smooth-wall computations, indicating that roughness may have augmented heating. The stagnation region also experienced heating that exceeded computational levels, but shock layer radiation does not fully explain the differences.

Edquist, Karl T.↗
NASA NTRS2013

Testing of DLR C/C-SiC for HIFiRE 8 Scramjet Combustor

Ceramic Matrix Composites (CMCs) have been proposed for hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a C/C and a C/C-SiC material system fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for HIFiRE 8, a joint Australia / AFRL hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kPa. Flat panels of the DLR C/C and the C/C-SiC were tested in the NASA Langley Direct Connect Rig (DCR) at Mach 5 and Mach 6 enthalpy for several minutes. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used in some of the tests to increase the surface temperature of the C/C-SiC panel for approximately 350degF. The final C/C-SiC panel was tested for 3 cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

Glass, David E.↗
NASA NTRS2013

Post-flight Analysis of Mars Science Laboratory Entry Aerothermal Environment and Thermal Protection System Response

The Mars Science Laboratory successfully landed on the Martian surface on August 5th, 2012. The rover was protected from the extreme heating environments of atmospheric entry by an ablative heatshield. This Phenolic Impregnated Carbon Ablator heatshield was instrumented with a suite of embedded thermocouples, isotherm sensors, and pressure transducers. The sensors monitored the in-depth ablator response, as well as the surface pressure at discrete locations throughout the hypersonic deceleration. This paper presents a comparison of the flight data with post-entry estimates. An assessment of the aerothermal environments, as well as the in-depth response of the heatshield material is made, and conclusions regarding the overall performance of the ablator at the suite locations are presented.

mars science laboratory↗
NASA NTRS2013

Aeroheating Measurement of Apollo Shaped Capsule with Boundary Layer Trip in the Free-piston Shock Tunnel HIEST

An aeroheating measurement test campaign of an Apollo capsule model with laminar and turbulent boundary layer was performed in the free-piston shock tunnel HIEST at JAXA Kakuda Space Center. A 250mm-diameter 6.4%-scaled Apollo CM capsule model made of SUS-304 stainless steel was applied in this study. To measure heat flux distribution, the model was equipped with 88 miniature co-axial Chromel-Constantan thermocouples on the heat shield surface of the model. In order to promote boundary layer transition, a boundary layer trip insert with 13 "pizza-box" isolated roughness elements, which have 1.27mm square, were placed at 17mm below of the model geometric center. Three boundary layer trip inserts with roughness height of k=0.3mm, 0.6mm and 0.8mm were used to identify the appropriate height to induce transition. Heat flux records with or without roughness elements were obtained for model angles of attack 28º under stagnation enthalpy between H(sub 0)=3.5MJ/kg to 21MJ/kg and stagnation pressure between P(sub 0)=14MPa to 60MPa. Under the condition above, Reynolds number based on the model diameter was varied from 0.2 to 1.3 million. With roughness elements, boundary layer became fully turbulent less than H(sub 0)=9MJ/kg condition. However, boundary layer was still laminar over H(sub 0)=13MJ/kg condition even with the highest roughness elements. An additional experiment was also performed to correct unexpected heat flux augmentation observed over H(sub 0)=9MJ/kg condition.

Hideyuki, TANNO↗
NASA NTRS2012

SPRITE: A TPS Test Bed for Ground and Flight

Engineers in the Entry Systems and Technology Division at NASA Ames Research Center developed a fully instrumented, small atmospheric entry probe called SPRITE (Small Probe Reentry Investigation for TPS Engineering). SPRITE, conceived as a flight test bed for thermal protection materials, was tested at full scale in an arc-jet facility so that the aerothermal environments the probe experiences over portions of its flight trajectory and in the arc-jet are similar. This ground-to-flight traceability enhances the ability of mission designers to evaluate margins needed in the design of thermal protection systems (TPS) of larger scale atmospheric entry vehicles. SPRITE is a 14-inch diameter, 45 deg. sphere-cone with a conical aftbody and designed for testing in the NASA Ames Aerodynamic Heating Facility (AHF). The probe is a two-part aluminum shell with PICA (phenolic impregnated carbon ablator) bonded on the forebody and LI-2200 (Shuttle tile material) bonded to the aftbody. Plugs with embedded thermocouples, similar to those installed in the heat shield of the Mars Science Laboratory (MSL), and a number of distributed sensors are integrated into the design. The data from these sensors are fed to an innovative, custom-designed data acquisition system also integrated with the test article. Two identical SPRITE models were built and successfully tested in late 2010-early 2011, and the concept is currently being modified to enable testing of conformable and/or flexible materials.

Ablation↗
NASA NTRS2014

Detailed Uncertainty Analysis of the ZEM-3 Measurement System

The measurement of Seebeck coefficient and electrical resistivity are critical to the investigation of all thermoelectric systems. Therefore, it stands that the measurement uncertainty must be well understood to report ZT values which are accurate and trustworthy. A detailed uncertainty analysis of the ZEM-3 measurement system has been performed. The uncertainty analysis calculates error in the electrical resistivity measurement as a result of sample geometry tolerance, probe geometry tolerance, statistical error, and multi-meter uncertainty. The uncertainty on Seebeck coefficient includes probe wire correction factors, statistical error, multi-meter uncertainty, and most importantly the cold-finger effect. The cold-finger effect plagues all potentiometric (four-probe) Seebeck measurement systems, as heat parasitically transfers through thermocouple probes. The effect leads to an asymmetric over-estimation of the Seebeck coefficient. A thermal finite element analysis allows for quantification of the phenomenon, and provides an estimate on the uncertainty of the Seebeck coefficient. The thermoelectric power factor has been found to have an uncertainty of +9-14 at high temperature and 9 near room temperature.

Seebeck effect↗
NASA NTRS2014

Uncertainty Analysis of Seebeck Coefficient and Electrical Resistivity Characterization

In order to provide a complete description of a materials thermoelectric power factor, in addition to the measured nominal value, an uncertainty interval is required. The uncertainty may contain sources of measurement error including systematic bias error and precision error of a statistical nature. The work focuses specifically on the popular ZEM-3 (Ulvac Technologies) measurement system, but the methods apply to any measurement system. The analysis accounts for sources of systematic error including sample preparation tolerance, measurement probe placement, thermocouple cold-finger effect, and measurement parameters; in addition to including uncertainty of a statistical nature. Complete uncertainty analysis of a measurement system allows for more reliable comparison of measurement data between laboratories.

Seebeck effect↗
NASA NTRS2014

An Experimental and Computational Study on Soot Formation in a Coflow Jet Flame Under Microgravity and Normal Gravity

Upon the completion of the Structure and Liftoff in Combustion Experiment (SLICE) in March 2012, a comprehensive and unique set of microgravity coflow diffusion flame data was obtained. This data covers a range of conditions from weak flames near extinction to strong, highly sooting flames, and enabled the study of gravitational effects on phenomena such as liftoff, blowout and soot formation. The microgravity experiment was carried out in the Microgravity Science Glovebox (MSG) on board the International Space Station (ISS), while the normal gravity experiment was performed at Yale utilizing a copy of the flight hardware. Computational simulations of microgravity and normal gravity flames were also carried out to facilitate understanding of the experimental observations. This paper focuses on the different sooting behaviors of CH4 coflow jet flames in microgravity and normal gravity. The unique set of data serves as an excellent test case for developing more accurate computational models.Experimentally, the flame shape and size, lift-off height, and soot temperature were determined from line-of-sight flame emission images taken with a color digital camera. Soot volume fraction was determined by performing an absolute light calibration using the incandescence from a flame-heated thermocouple. Computationally, the MC-Smooth vorticity-velocity formulation was employed to describe the chemically reacting flow, and the soot evolution was modeled by the sectional aerosol equations. The governing equations and boundary conditions were discretized on an axisymmetric computational domain by finite differences, and the resulting system of fully coupled, highly nonlinear equations was solved by a damped, modified Newtons method. The microgravity sooting flames were found to have lower soot temperatures and higher volume fraction than their normal gravity counterparts. The soot distribution tends to shift from the centerline of the flame to the wings from normal gravity to microgravity.

microgravity↗
NASA NTRS2014

Testing of DLR C/C-SiC and C/C for HIFiRE 8 Scramjet Combustor

Ceramic Matrix Composites (CMCs) have been proposed for use as lightweight hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a carbon/carbon (C/C) and a carbon/carbon-silicon carbide (C/C-SiC) material fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for Hypersonic International Flight Research Experimentation (HIFiRE) 8, a joint Australia / Air Force Research Laboratory hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kilopascals. Flat panels of the DLR C/C and C/C-SiC materials were installed downstream of a hydrogen-fueled, dual-mode scramjet combustor and tested for several minutes at conditions simulating flight at Mach 5 and Mach 6. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used during some of the tests to reduce heat loss from the back surface and thus increase the heated surface temperature of the C/C-SiC panel approximately 177 C (350 F). The final C/C-SiC panel was tested for three cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

Glass, David E.↗
NASA NTRS2014

Validation Ice Crystal Icing Engine Test in the Propulsion Systems Laboratory at NASA Glenn Research Center

The Propulsion Systems Laboratory (PSL) is an existing altitude simulation jet engine test facility located at NASA Glenn Research Center in Clevleand, OH. It was modified in 2012 with the integration of an ice crystal cloud generation system. This paper documents the inaugural ice crystal cloud test in PSLthe first ever full scale, high altitude ice crystal cloud turbofan engine test to be conducted in a ground based facility. The test article was a Lycoming ALF502-R5 high bypass turbofan engine, serial number LF01. The objectives of the test were to validate the PSL ice crystal cloud calibration and engine testing methodologies by demonstrating the capability to calibrate and duplicate known flight test events that occurred on the same LF01 engine and to generate engine data to support fundamental and computational research to investigate and better understand the physics of ice crystal icing in a turbofan engine environment while duplicating known revenue service events and conducting test points while varying facility and engine parameters. During PSL calibration testing it was discovered than heated probes installed through tunnel sidewalls experienced ice buildup aft of their location due to ice crystals impinging upon them, melting and running back. Filtered city water was used in the cloud generation nozzle system to provide ice crystal nucleation sites. This resulted in mineralization forming on flow path hardware that led to a chronic degradation of performance during the month long test. Lacking internal flow path cameras, the response of thermocouples along the flow path was interpreted as ice building up. Using this interpretation, a strong correlation between total water content (TWC) and a weaker correlation between median volumetric diameter (MVD) of the ice crystal cloud and the rate of ice buildup along the instrumented flow path was identified. For this test article the engine anti-ice system was required to be turned on before ice crystal icing would occur. The ice crystal icing event, an uncommanded reduction in thrust, was able to be turned on and off by manipulating cloud TWC. A flight test point where no ice crystal icing event occurred was also duplicated in PSL. Physics based computational tools were successfully used to predict tunnel settings to induce ice buildup along the low pressure compression system flow path for several test points at incrementally lower altitudes, demonstrating that development of ice crystal icing scaling laws is potentially feasible. Analysis of PSL test data showed that uncommanded reduction in thrust occurs during ice crystal cloud on operation prior to fan speed reduction. This supports previous findings that the reduction of thrust for this test article is due to ice buildup leading to a restricted airflow from either physical or aerodynamic blockage in the engine core flow path.

Ice Crystal Icing↗
NASA NTRS2014

Validation Ice Crystal Icing Engine Test in the Propulsion Systems Laboratory at NASA Glenn Research Center

The Propulsion Systems Laboratory (PSL) is an existing altitude simulation jet engine test facility located at NASA Glenn Research Center in Cleveland, OH. It was modified in 2012 with the integration of an ice crystal cloud generation system. This paper documents the inaugural ice crystal cloud test in PSL--the first ever full scale, high altitude ice crystal cloud turbofan engine test to be conducted in a ground based facility. The test article was a Lycoming ALF502-R5 high bypass turbofan engine, serial number LF01. The objectives of the test were to validate the PSL ice crystal cloud calibration and engine testing methodologies by demonstrating the capability to calibrate and duplicate known flight test events that occurred on the same LF01 engine and to generate engine data to support fundamental and computational research to investigate and better understand the physics of ice crystal icing in a turbofan engine environment while duplicating known revenue service events and conducting test points while varying facility and engine parameters. During PSL calibration testing it was discovered than heated probes installed through tunnel sidewalls experienced ice buildup aft of their location due to ice crystals impinging upon them, melting and running back. Filtered city water was used in the cloud generation nozzle system to provide ice crystal nucleation sites. This resulted in mineralization forming on flow path hardware that led to a chronic degradation of performance during the month long test. Lacking internal flow path cameras, the response of thermocouples along the flow path was interpreted as ice building up. Using this interpretation, a strong correlation between total water content (TWC) and a weaker correlation between median volumetric diameter (MVD) of the ice crystal cloud and the rate of ice buildup along the instrumented flow path was identified. For this test article the engine anti-ice system was required to be turned on before ice crystal icing would occur. The ice crystal icing event, an uncommanded reduction in thrust, was able to be turned on and off by manipulating cloud TWC. A flight test point where no ice crystal icing event occurred was also duplicated in PSL. Physics based computational tools were successfully used to predict tunnel settings to induce ice buildup along the low pressure compression system flow path for several test points at incrementally lower altitudes, demonstrating that development of ice crystal icing scaling laws is potentially feasible. Analysis of PSL test data showed that uncommanded reduction in thrust occurs during ice crystal cloud on operation prior to fan speed reduction. This supports previous findings that the reduction of thrust for this test article is due to ice buildup leading to a restricted airflow from either physical or aerodynamic blockage in the engine core flow path.

Turbofan Engine↗
NASA NTRS2014

Advances in Current Rating Techniques for Flexible Printed Circuits

Twist Capsule Assemblies are power transfer devices commonly used in spacecraft mechanisms that require electrical signals to be passed across a rotating interface. Flexible printed circuits (flex tapes, see Figure 2) are used to carry the electrical signals in these devices. Determining the current rating for a given trace (conductor) size can be challenging. Because of the thermal conditions present in this environment the most appropriate approach is to assume that the only means by which heat is removed from the trace is thru the conductor itself, so that when the flex tape is long the temperature rise in the trace can be extreme. While this technique represents a worst-case thermal situation that yields conservative current ratings, this conservatism may lead to overly cautious designs when not all traces are used at their full rated capacity. A better understanding of how individual traces behave when they are not all in use is the goal of this research. In the testing done in support of this paper, a representative flex tape used for a flight Solar Array Drive Assembly (SADA) application was tested by energizing individual traces (conductors in the tape) in a vacuum chamber and the temperatures of the tape measured using both fine-gauge thermocouples and infrared thermographic imaging. We find that traditional derating schemes used for bundles of wires do not apply for the configuration tested. We also determine that single active traces located in the center of a flex tape operate at lower temperatures than those on the outside edges.

Hayes, Ron↗
NASA NTRS2014

Long Hole Film Cooling Dataset for CFD Development - Flow and Film Effectiveness

An experiment investigating flow and heat transfer of long (length to diameter ratio of 18) cylindrical film cooling holes has been completed. In this paper, the thermal field in the flow and on the surface of the film cooled flat plate is presented for nominal freestream turbulence intensities of 1.5 and 8 percent. The holes are inclined at 30 deg above the downstream direction, injecting chilled air of density ratio 1.0 onto the surface of a flat plate. The diameter of the hole is 0.75 in. (approx. 0.02 m) with center to center spacing (pitch) of 3 hole diameters. Coolant was injected into the mainstream flow at nominal blowing ratios of 0.5, 1.0, 1.5, and 2.0. The Reynolds number of the freestream was approximately 11,000 based on hole diameter. Thermocouple surveys were used to characterize the thermal field. Infrared thermography was used to determine the adiabatic film effectiveness on the plate. Hotwire anemometry was used to provide flowfield physics and turbulence measurements. The results are compared to existing data in the literature. The aim of this work is to produce a benchmark dataset for Computational Fluid Dynamics (CFD) development to eliminate the effects of hole length to diameter ratio and to improve resolution in the near-hole region. In this report, a Time Filtered Navier Stokes (TFNS), also known as Partially Resolved Navier Stokes (PRNS), method that was implemented in the Glenn-HT code is used to model coolant-mainstream interaction. This method is a high fidelity unsteady method that aims to represent large scale flow features and mixing more accurately.

Shyam, Vikram↗
NASA NTRS2009

Solar Thermal Vacuum Test of Deployable Astromesh Reflector

On September 10, 2008, a 36-hour Solar Thermal Vacuum Test of a 5m deployable mesh reflector was completed in JPL's 25' Space Simulator by the Advanced Deployable Structures Group at JPL. The testing was performed under NASA's Innovative Partnership Program (IPP) as a risk reduction effort for two JPL Decadal Survey Missions: DESDynI and SMAP. The 5.0 m aperture Astromesh reflector was provided by Northrop Grumman Aerospace Systems (NGAS) Astro Aerospace, our IPP industry partner. The testing utilized a state-of-the-art photogrammetry system to measure deformation of the reflector under LN2 cold soak, 0.25 Earth sun, 0.5 sun and 1.0 sun. An intricate network of thermocouples (approximately 200 in total) was used to determine the localized temperature across the mesh as well as on the perimeter truss of the reflector. Half of the reflector was in a fixed shadow to maximize thermal gradients. A mobility system was built for remotely actuating the cryo-vacuum capable photogrammetry camera around the circumference of the Solar Simulator. Photogrammetric resolution of 0.025 mm RMS (0.001") was achieved over the entire 5 meter aperture for each test case. The data will be used for thermo-elastic model correlation and validation, which will benefit the planned Earth Science Missions.

Photogrammetry↗