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NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗
NASA NTRS1973

Gas flow analysis during thermal vacuum test of a spacecraft.

The pressures indicated by two tubulated ionization gages, one pointing to a spinning spacecraft undergoing thermal vacuum test and the other the walls of the chamber, have been used in a computer program to calculate important parameters of flow kinetics in the vacuum chamber. These parameters calculated as a function of time are: the self-contamination of the spacecraft (defined as the return of outgassed molecules on its critical surfaces either in orbit or while undergoing vacuum test); the spacecraft outgassing including leaks from sealed compartments; and the gas pumping performance of the vacuum chamber. The test indicated the feasibility of this type of evaluation and the improvements in instrumentations and arrangements needed for future tests.

Scialdone, J. J.↗
NASA NTRS1988

Development of a two axis motion simulation system for thermal/vacuum satellite testing

A two-axis motion simulation system for thermal vacuum testing of large satellites in a space simulation chamber was developed. Satellites as large as 3000 kilograms with a 4-meter diameter and a 5-meter length can be tested. This motion simulator (MS) incorporates several unique features which result in a less complicated design with improved performance when compared to previous satellite motion simulators. The design of the simulator is discussed in detail.

Henderson, David↗
NASA NTRS1993

CFM technologies for space transportation: Multipurpose hydrogen testbed system definition and tank procurement

The development of a test bed tank and system for evaluating cryogenic fluid management technologies in a simulated upper stage liquid hydrogen tank is covered. The tank is 10 ft long and is 10 ft in diameter, and is an ASME certified tank constructed of 5083 aluminum. The tank is insulated with a combination of sprayed on foam insulation, covered by 45 layers of double aluminized mylar separated by dacron net. The mylar is applied by a continuous wrap system adapted from commercial applications, and incorporates variable spacing between the mylar to provide more space between those layers having a high delta temperature, which minimizes heat leak. It also incorporates a unique venting system which uses fewer large holes in the mylar rather than the multitude of small holes used conventionally. This significantly reduces radiation heat transfer. The test bed consists of an existing vacuum chamber at MSFC, the test bed tank and its thermal control system, and a thermal shroud (which may be heated) surrounding the tank. Provisions are made in the tank and chamber for inclusion of a variety of cryogenic fluid management experiments.

Fox, E. C.↗
NASA NTRS1994

Refurbishment of a 39 foot thermal vacuum chamber

The 39' thermal vacuum chamber at Space Systems/Loral has been used to test numerous spacecraft including those of the GOES, Intelsat, Insat, Superbird, N-Star, NATO and other programs. Ten years ago, the aluminum LN2 shroud experienced serious fatigue failures in the field welded jumper tubing, effectively shutting down the chamber for vacuum testing. The problem was repaired at the time, but new failures began to reappear a few months ago and are now occurring at a rate that suggests that the shroud may again become inoperable. Consequently, Space Systems/Loral is spending in excess of $6 million to replace the shroud and the existing LN2 equipment with a new, state of the art cryogenic system. In May, 1994, a contract was awarded to remove the existing shroud and LN2 pumping system and replace it with a gravity fed shroud and distribution system. Included in the contract are eight skid mounted gaseous nitrogen pumping systems capable of controlling shroud zone temperatures between +150 C and -180 C. The project is scheduled to be completed in April 1995.

Edwards, Arthur A.↗
NASA NTRS1994

New potentials of NIICHIMMASH's thermal vacuum facilities

The potentialities of existing test facilities as to simulating space environment governing factors for spacecraft successful development thermal vacuum testing are analyzed, ways of modernizing existing test facilities and specific proposals on their redesign are considered. The problem of spacecraft (S/C) ground development in simulated external environments, the solution of which started more than 30 years ago, has not lost its urgency today. Stringent requirements on S/C active lifetime under space conditions, module large dimensions, great number of extension elements and complicated mode of their interaction in long mission do not allow S/C designers to abandon ground tests. S/C thermal modes development is a combination of calculations, thermal vacuum tests and actions on improving S/C design and its thermal control system. Traditionally, tests are carried out by stages from component and end unit level verifications to complex tests of modules and S/C as a whole. In our opinion, sufficient correctness of calculated models and experience gained in organizations designing space systems allow to reduce cost and time of autonomous tests. Unfortunately, this is not true for complex (integrated) thermal vacuum tests. More than that, their recent programs include tasks of verifying other (than thermal control system) systems if S/C for operation under space simulated conditions. The outlined circumstances are the main reason for critical review of the potentialities of the existing test base, and of NIICHIMMASH's two large thermal vacuum chambers, first of all. The reasons for and ways of enlargement of these facilities potentially are analyzed and the results attained are described.

Afanassiev, N. A.↗
NASA NTRS2014

Temperature Knowledge and Model Correlation for the Soil Moisture Active and Passive (SMAP) Reflector Mesh

NASA's Earth observing Soil Moisture Active & Passive (SMAP) Mission is scheduled to launch in November 2014 into a 685 km near-polar, sun synchronous orbit. SMAP will provide comprehensive global mapping measurements of soil moisture and freeze/thaw state in order to enhance understanding of the processes that link the water, energy, and carbon cycles. The primary objectives of SMAP are to improve worldwide weather and flood forecasting, enhance climate prediction, and refine drought and agriculture monitoring during its 3 year mission. The SMAP instrument architecture incorporates an L-band radar and an L-band radiometer which share a common feed horn and parabolic mesh reflector. The instrument rotates about the nadir axis at approximately 15 rpm, thereby providing a conically scanning wide swath antenna beam that is capable of achieving global coverage within 3 days. In order to make the necessary precise surface emission measurements from space, a temperature knowledge of 60 deg C for the mesh reflector is required. In order to show compliance, a thermal vacuum test was conducted using a portable solar simulator to illuminate a non flight, but flight-like test article through the quartz window of the vacuum chamber. The molybdenum wire of the antenna mesh is too fine to accommodate thermal sensors for direct temperature measurements. Instead, the mesh temperature was inferred from resistance measurements made during the test. The test article was rotated to five separate angles between 10 deg and 90 deg via chamber breaks to simulate the maximum expected on-orbit solar loading during the mission. The resistance measurements were converted to temperature via a resistance versus temperature calibration plot that was constructed from data collected in a separate calibration test. A simple thermal model of two different representations of the mesh (plate and torus) was created to correlate the mesh temperature predictions to within 60 deg C. The on-orbit mesh temperature will be predicted using the correlated analytical thermal model since direct measurements from in-situ flight thermal sensors are not possible.

Mikhaylov, Rebecca↗
NASA NTRSDate not supplied

Solar White Thermal Coating for Cryogenic Propulsion Systems

NASA is currently conducting research into the potential of storing cryogenic fluid in low Earth orbit (LEO). Having cryogenic propellant readily available for high-performance propulsion systems can be very beneficial for deep space missions in the near future. One of the key challenges to storing cryogenic fluid in LEO is minimizing boil-off. To address the challenge, NASA is evaluating new concepts in thermal insulation. One recent experimental study evaluated the feasibility of using Yttrium Oxide (Y2O3), formed into tiles or spray coating that can potentially be used as a thermal coating for cryogenic propellant storage applications in deep space. Due to its temperature and wavelength dependent optical properties, this “solar white” material can reflect a vast majority of the Sun’s radiative energy while having a very high infrared emissivity for rejecting heat to deep space. As a part of the material development and proof of concept testing, multiple tests have been run at KSC and GRC to demonstrate the performance of the material. At GRC, the tests were run using the Deep Space Solar Simulator (DS3) which contains a thermal vacuum chamber in which the solar white sample was exposed to a deep space environment (< 10 K, optically dense walls) while under full illumination via solar lamp. In order to improve the use of the test results and apply them to spacecraft, there is a need to be able to model the material properties within NASA’s standard thermal modeling tools. As such, it was set out to verify a thermal model of one of the experiments using Thermal Desktop with Sinda. The thermal model of the DS3 test setup includes the solar white sample along with a solar lamp simulator capable of outputting heat at varying wavelengths. The model was developed in order to validate the test results and also to help predict results that will be obtained in future tests. This paper will review the modeling methods and thermal analysis results for various test cases that were run within the DS3 facility.

Cryogenic↗
NASA NTRS2015

Impact of Drilling Operations on Lunar Volatiles Capture: Thermal Vacuum Tests

In Situ Resource Utilization (ISRU) enables future planetary exploration by using local resources to supply mission consumables. This idea of 'living off the land' has the potential to reduce mission cost and risk. On the moon, water has been identified as a potential resource (for life support or propellant) at the lunar poles, where it exists as ice in the subsurface. However, the depth and content of this resource has yet to be confirmed on the ground; only remote detection data exists. The upcoming Resource Prospector mission (RP) will 'ground-truth' the water using a rover, drill, and the RESOLVE science package. As the 2020 planned mission date nears, component level hardware is being tested in relevant lunar conditions (thermal vacuum). In August 2014 a series of drilling tests were performed using the Honeybee Robotics Lunar Prospecting Drill inside a 'dirty' thermal vacuum chamber at the NASA Glenn Research Center. The drill used a unique auger design to capture and retain the lunar regolith simulant. The goal of these tests was to investigate volatiles (water) loss during drilling and sample transfer to a sample crucible in order to validate this regolith sampling method. Twelve soil samples were captured over the course of two tests at pressures of 10(exp-5) Torr and ambient temperatures between -80C to -20C. Each sample was obtained from a depth of 40 cm to 50 cm within a cryogenically frozen bed of NU-LHT-3M lunar regolith simulant doped with 5 wt% water. Upon acquisition, each sample was transferred and hermetically sealed inside a crucible. The samples were later baked out to determine water wt% and in turn volatile loss by following ASTM standard practices. Of the twelve tests, four sealed properly and lost an average of 30% of their available water during drilling and transfer. The variability in the results correlated well with ambient temperature (lower the temperature lower volatiles loss) and the trend agreed with the sublimation rates for the same temperature. Moisture retention also correlated with quantity of sample: a larger amount of material resulted in less water loss. The drilling process took an average of 10 minutes to capture and transfer each sample. The drilling power was approximately 20 Watt with a Weight on Bit of approximately 30 N. The bit temperature indicated little heat input into formation during the drilling process.

drilling↗
NASA NTRS2012

Thermal and Alignment Analysis of the Instrument-Level ATLAS Thermal Vacuum Test

This paper describes the thermal analysis and test design performed in preparation for the ATLAS thermal vacuum test. NASA's Advanced Topographic Laser Altimeter System (ATLAS) will be flown as the sole instrument aboard the Ice, Cloud, and land Elevation Satellite-2 (ICESat-2). It will be used to take measurements of topography and ice thickness for Arctic and Antarctic regions, providing crucial data used to predict future changes in worldwide sea levels. Due to the precise measurements ATLAS is taking, the laser altimeter has very tight pointing requirements. Therefore, the instrument is very sensitive to temperature-induced thermal distortions. For this reason, it is necessary to perform a Structural, Thermal, Optical Performance (STOP) analysis not only for flight, but also to ensure performance requirements can be operationally met during instrument-level thermal vacuum testing. This paper describes the thermal model created for the chamber setup, which was used to generate inputs for the environmental STOP analysis. This paper also presents the results of the STOP analysis, which indicate that the test predictions adequately replicate the thermal distortions predicted for flight. This is a new application of an existing process, as STOP analyses are generally performed to predict flight behavior only. Another novel aspect of this test is that it presents the opportunity to verify pointing results of a STOP model, which is not generally done. It is possible in this case, however, because the actual pointing will be measured using flight hardware during thermal vacuum testing and can be compared to STOP predictions.

Bradshaw, Heather↗
NASA NTRS2013

Creating the Deep Space Environment for Testing the James Webb Space Telescope at NASA Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft.) in diameter and 36.6 m (120 ft.) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. It was originally designed and built in the mid 1960 s to test the Apollo Command and Service Module and several manned tests were conducted on that spacecraft, contributing to the success of the program. The chamber has been used since that time to test spacecraft active thermal control systems, Shuttle DTO, DOD, and ESA hardware in simulated Low Earth Orbit (LEO) conditions. NASA is now moving from LEO towards exploration of locations with environments approaching those of deep space. Therefore, Chamber A has undergone major modifications to enable it to simulate these deeper space environments. Environmental requirements were driven, and modifications were funded by the James Webb Space Telescope program, and this telescope, which will orbit Solar/Earth L2, will be the first test article to benefit from the chamber s new capabilities. To accommodate JWST, the Chamber A high vacuum system has been modernized, additional LN2 shrouds have been installed, the liquid nitrogen system has been modified to minimize dependency on electrical power and increase its reliability, a new helium shroud/refrigeration system has been installed to create a colder more stable and uniform heat sink, and the controls have been updated to increase the level of automation and improve operator interfaces. Testing of these major modifications was conducted in August of 2012 and this initial test was very successful, with all major systems exceeding their performance requirements. This paper will outline the changes in overall environmental requirements, discuss the technical design data that was used in the decisions leading to the extensive modifications, and describe the new capabilities of the chamber.

Homan, Jonathan L.↗
NASA NTRS2012

Creating the Deep Space Environment for Testing the James Webb Space Telescope at the Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft) in diameter and 36.6 m (120 ft) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. It was originally designed and built in the mid 1960's to test the Apollo Command and Service Module and several manned tests were conducted on that spacecraft, contributing to the success of the program. The chamber has been used since that time to test spacecraft active thermal control systems, Shuttle DTO, DOD, and ESA hardware in simulated Low Earth Orbit (LEO) conditions. NASA is now moving from LEO towards exploration of locations with environments approaching those of deep space. Therefore, Chamber A has undergone major modifications to enable it to simulate these deeper space environments. Environmental requirements were driven, and the modifications were funded, by the James Webb Space Telescope program, and this telescope which will orbit Solar/Earth L2, will be the first test article to benefit from the chamber s new capabilities. To accommodate JWST, the Chamber A high vacuum system has been modernized, additional LN2 shrouds have been installed, the liquid nitrogen system has been modified to remove dependency on electrical power and increase its reliability, a new helium shroud/refrigeration system has been installed to create a colder more stable and uniform heat sink and, the controls have been updated to increase the level of automation and improve operator interfaces. Testing of these major modifications was conducted in August 2012 and this initial test was very successful, with all major systems exceeding their performance requirements. This paper will outline the changes in the overall environmental requirements, discuss the technical design data that was used in the decisions leading to the extensive modifications, and describe the new capabilities of the chamber.

Homan, Jonathan L.↗
NASA NTRS2026

xEMU Spacesuit Boot Lunar South Pole Thermal Vacuum Testing

The Exploration Extravehicular Mobility Unit (xEMU) spacesuit lunar boot was evaluated in the Jet Propulsion Lab (JPL) CITADEL thermal vacuum chamber at lunar South Pole conditions. The JPL CITADEL chamber provides a radiation environment and contact plate temperature of 48K, simulating the temperature of a lunar Permanently Shadowed Region (PSR) at the South Pole of the Moon. The main objectives of the test series were: to evaluate the risk lunar PSRs pose to advanced spacesuit boots, validate the xEMU lunar boot thermal model, and develop a test methodology to evaluate lunar spacesuit boots. To simulate the thermal effects of a crewmember’s foot inside the boot, a thermal manikin foot was developed to inject heat into the spacesuit boot. Thermal data was collected through a variety of sensors including thermocouples, thermistors, heat flux sensors, and infrared cameras. A test matrix was developed to investigate the variables of foot contact force, simulated skin temperature, external contact plate temperature, and ventilation flowrate. The JPL CITADEL chamber includes a loadlock (airlock) that afforded the ability to simulate the transition of the boot from a thermally neutral IVA condition to a cold EVA condition within seconds. This paper will document the test design, test objectives, success criteria, test support equipment, spacesuit boot to chamber interfaces, and configuration of the test article. Additionally, this paper will document the high-level results of the testing. More detailed test results will be presented in a different paper.

TVAC↗
NASA NTRS2006

Solar Sails

The Solar Sail Propulsion investment area has been one of the three highest priorities within the In-Space Propulsion Technology (ISPT) Project. In the fall of 2003, the NASA Headquarters' Science Mission Directorate provided funding and direction to mature the technology as far as possible through ground research and development from TRL 3 to 6 in three years. A group of experts from government, industry, and academia convened in Huntsville, Alabama to define technology gaps between what was needed for science missions to the inner solar system and the current state of the art in ultra1ightweight materials and gossamer structure design. This activity set the roadmap for development. The centerpiece of the development would be the ground demonstration of scalable solar sail systems including masts, sails, deployment mechanisms, and attitude control hardware and software. In addition, new materials would be subjected to anticipated space environments to quantify effects and assure mission life. Also, because solar sails are huge structures, and it is not feasible to validate the technology by ground test at full scale, a multi-discipline effort was established to develop highly reliable analytical models to serve as mission assurance evidence in future flight program decision-making. Two separate contractor teams were chosen to develop the SSP System Ground Demonstrator (SGD). After a three month conceptual mission/system design phase, the teams developed a ten meter diameter pathfinder set of hardware and subjected it to thermal vacuum tests to compare analytically predicted structural behavior with measured characteristics. This process developed manufacturing and handling techniques and refined the basic design. In 2005, both contractor teams delivered 20 meter, four quadrant sail systems to the largest thermal vacuum chamber in the world in Plum Brook, Ohio, and repeated the tests. Also demonstrated was the deployment and articulation of attitude control mechanisms. In conjunction with these tests, the stowed sails were subjected to launch vibration and ascent vent tests. Other investments studied radiation effects on the solar sail materials, investigated spacecraft charging issues, developed shape measuring techniques and instruments, produced advanced trajectory modeling capabilities, and identified and resolved gossamer structure dynamics issues. Technology validation flight and application to a He1iophysics science mission is on the horizon.

Young, Roy↗
NASA NTRS2013

Creating the Deep Space Environment for Testing the James Webb Space Telescope (JWST) at NASA Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft) in diameter and 36.6 m (120 ft) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. It was originally designed and built in the mid 1960 s to test the Apollo Command and Service Module and several manned tests were conducted on that spacecraft, contributing to the success of the program. The chamber has been used since that time to test spacecraft active thermal control systems, Shuttle DTO, DOD, and ESA hardware in simulated Low Earth Orbit (LEO) conditions. NASA is now moving from LEO towards exploration of locations with environments approaching those of deep space. Therefore, Chamber A has undergone major modifications to enable it to simulate these deeper space environments. Environmental requirements were driven, and modifications were funded by the James Webb Space Telescope program, and this telescope which will orbit Solar/Earth L2, will be the first test article to benefit from the chamber s new capabilities. To accommodate JWST, the Chamber A high vacuum system has been modernized, additional LN2 shrouds have been installed, the liquid nitrogen system has been modified to remove dependency on electrical power and increase its reliability, a new helium shroud/refrigeration system has been installed to create a colder more stable and uniform heat sink, and the controls have been updated to increase the level of automation and improve operator interfaces. Testing of these major modifications was conducted in August of 2012 and this initial test was very successful, with all major systems exceeding their performance requirements. This paper will outline the changes in overall environmental requirements, discuss the technical design data that was used in the decisions leading to the extensive modifications, and describe the new capabilities of the chamber.

Homan, Jonathan L.↗
NASA NTRS1990

Thermal balance testing of the Anik-E spacecraft

Spar Aerospace, Ltd. of Canada has recently completed the first of two infrared thermal balance/thermal vacuum tests on Canada's next generation of telecommunication satellites: Anik-E. Anik-E is a two spacecraft program using a three axis stabilized GE 5000 series bus as the vehicle for its payload. The infrared test setup consists of the mechanical hardware required to support the spacecraft within the thermal vacuum chamber and the infrared rig which supports the spacecraft. Spar developed a software package which calculated the lamp positions to provide the necessary uniform heat flux, based on the required heat inputs on the various isothermal zones of the spacecraft for the different phases of the tests. Infrared heat sensors, radiometers, were installed on the spacecraft surfaces. These were connected to the facility data processing system and continually monitored the flux absorbed by the various spacecraft surfaces. The test set up had distinct advantages. It enabled testing to progress directly from the thermal balance to thermal vacuum. Waveguide runs were installed for the communications portion of the testing prior to pumpdown. This precluded the requirement to return to ambient to reconfigure the setup.

Golob, Alex↗
NASA NTRS2016

Creating the Thermal Environment for Safely Testing the James Webb Space Telescope at the Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft) in diameter and 36.6 m (120 ft) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. The chamber was originally built to support testing of the Apollo Service and Command Module for lunar missions, but underwent major modifications to be able to test the James Webb Space Telescope in a simulated deep space environment. To date seven tests have been performed in preparation of testing the flight optics for the James Webb Space Telescope (JWST). Each test has had a uniquie thermal profile and set of thermal requirements for cooling down and warming up, controlling contamination, and releasing condensed air. These range from temperatures from 335K to 15K, with tight uniformity and controllability for maintining thermal stability and pressure control. One unique requirement for two test was structurally proof loading hardware by creating thermal gradients at specific temperatures. This paper will discuss the thermal requirements and goals of the tests, the original requirements of the chamber thermal systems for planned operation, and how the new requirements were met by the team using the hardware, system flexiblilty, and engineering creativity. It will also discuss the mistakes and successes to meet the unique goals, especially when meeting the thermal proof load.

Homan, Jonathan L.↗