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All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Experimental studies of transpiration cooling with shock interaction in hypersonic flow, part B

This report describes the result of experimental studies conducted to examine the effects of the impingement of an oblique shock on the flowfield and surface characteristics of a transpiration-cooled wall in turbulent hypersonic flow. The principal objective of this work was to determine whether the interaction between the oblique shock and the low-momentum region of the transpiration-cooled boundary layer created a highly distorted flowfield and resulted in a significant reduction in the cooling effectiveness of the transpiration-cooled surface. As a part of this program, we also sought to determine the effectiveness of transpiration cooling with nitrogen and helium injectants for a wide range of blowing rates under constant-pressure conditions in the absence of shock interaction. This experimental program was conducted in the Calspan 48-Inch Shock Tunnel at nominal Mach numbers of 6 and 8, for a Reynolds number of 7.5 x 10(exp 6). For these test conditions, we obtained fully turbulent boundary layers upstream of the interaction regions over the transpiration-cooled segment of the flat plate. The experimental program was conducted in two phases. In the first phase, we examined the effects of mass-addition level and coolant properties on the cooling effectiveness of transpiration-cooled surfaces in the absence of shock interaction. In the second phase of the program, we examined the effects of oblique shock impingement on the flowfield and surface characteristics of a transpiration-cooled surface. The studies were conducted for a range of shock strengths with nitrogen and helium coolants to examine how the distribution of heat transfer and pressure and the characteristics of the flowfield in the interaction region varied with shock strength and the level of mass addition from the transpiration-cooled section of the model. The effects of the distribution of the blowing rate along the interaction regions were also examined for a range of blowing rates through the transpiration-cooled panels. The regions of shockwave/boundary layer interaction examined in these studies were induced by oblique shocks generated with a sharp, flat plate, inclined to the freestream at angles of 5 degrees, 7.5 degrees, and 10 degrees. It was found that, in the absence of an incident shock, transpiration cooling was a very effective method for reducing both the heat transfer and the skin friction loads on the surface. The helium coolant was found to be significantly more effective than nitrogen, because of its low molecular weight and high specific heat. The studies of shock-wave/transpiration-cooled surface interaction demonstrated that the interaction region between the incident shock and the low-momentum transpiration-cooled boundary layer did not result in a significant increase in the size of attached or separated interaction regions, and did not result in significant flowfield distortions above the interaction region. The increase in heating downstream of the shock-impingement point could easily be reduced to the values without shock impingement by a relatively small increase in the transpiration cooling in this region. Surprisingly, this increase in cooling rate did not result in a significant increase in size of the region ahead of the incident shock or create a significantly enlarged interaction region with a resultant increase in the distortion level in the inviscid flow. Thus, transpiration cooling appears to be a very effective technique to cool the internal surfaces of scramjet engines, where shocks in the engine would induce large local increases in wall heating and create viscous/inviscid interactions that could significantly disturb the smooth flow through the combustor. However, if hydrogen is used as the coolant, burning upstream of shock impingement might result in localized hot spots. Clearly, further research is needed in this area.

Holden, Michael S.↗
NASA NTRS2003

A Compact, Continuous Adiabatic Demagnetization Refrigerator with High Heat Sink Temperature

In the continuous adiabatic demagnetization refrigerator (ADR), the existence of a constant temperature stage attached to the load breaks the link between the requirements of the load (usually a detector array) and the operation of the ADR. This allows the ADR to be cycled much faster, which yields more than an order of magnitude improvement in cooling power density over single-shot ADRs. Recent effort has focused on developing compact, efficient higher temperature stages. An important part of this work has been the development of passive gas-gap heat switches that transition (from conductive to insulating) at temperatures around 1 K and 4 K without the use of an actively heated getter. We have found that by carefully adjusting available surface area and the number of He-3 monolayers, gas-gap switches can be made to operate passively. Passive operation greatly reduces switching time and eliminates an important parasitic heat load. The current four stage ADR provides 6 micro W of cooling at 50 mK (21 micro W at 100 mK) and weighs less than 8 kg. It operates from a 4.2 K heat sink, which can be provided by an unpumped He bath or many commercially available mechanical cryocoolers. Reduction in critical current with temperature in our fourth stage NbTi magnet presently limits the maximum temperature of our system to approx. 5 K. We are developing compact, low-current Nb3Sn magnets that will raise the maximum heat sink temperature to over 10 K.

Shirron, P. J.↗
NASA NTRS1974

Shuttle active thermal control system development testing. Volume 1: Overall summary

A summary is given of a series of thermal vacuum tests designed to support the development of the orbiter active thermal control system (ATCS) and included testing of a wide heat load range modular radiator system (MRS) configured to the March 1973 orbiter baseline system, a candidate weight reducing radiator/water cooling system, and a smaller radiator system with a high performance radiator coating. The tests verified the performance of the baseline system and obtained detailed design information for application of a wide heat load range modular radiator system to the orbiter. The two candidate ATCS weight reducing designs have undergone extensive concept verification testing and their system operating characteristics have been determined in sufficient detail for application to the orbiter. Design information has been obtained for an integrated radiator/water cooling system that provides for vehicle heat rejection as well as water management of the excess fuel cell water. Processing techniques have been developed and verified for the application of a high performance thermal control coating to large radiator areas subjected to a temperature range of -280 F to +160 F.

Howell, H. R.↗
NASA NTRS2017

Feasibility of Passive Cryogenic Cooling for Solar Powered Outer Planetary Missions

Spacescience instruments with cooled detectors require innovative thermal cooling solutions to meet science objectives. Detector sensitivity increases with decreasing temperature and low optics temperatures are often needed to reduce background photon noise. As the detector spectral range coverage increases from the visible to far infrared also requires lower detector temperatures. Increasing demands on detector performance lead to larger format detectors and higher refresh rates resulting in significant increases in power dissipation. Passive coolers rely on emissive power of radiating surfaces to reject heat to space. As the operating temperature requirements of detectors and optics decreases, the ability to reject heat to space becomes increasingly more difficult. Reducing both cooler internal parasitic and external environmental heat loads and maximizing the passive cooler field of view to space will enhance performance. While instrument heat loads and passive cooler parasitic heat loads are controlled by instrument designers, the external environmental heat loads and cooler views to space are governed by spacecraft and mission designers. Solar powered planetary missions require large arrays to generate sufficient power for spacecraft subsystems and payloads. Two or more solar array wings with cell coverage of the order of 40-80 m2 are often needed to generate sufficient power at 3-6 AU. These large arrays are typically symmetrically configured and can extend tens of meters. The arrays along with spacecraft attitude requirements near the target planetary bodies pose significant challenges for passive cooling at large AU. It is very difficult to provide a clear field of view to space for the cooler with large articulating arrays while keeping solar loads from impinging the cooler and meeting the spacecraft attitude science pointing requirements. This is counter intuitive because of the decreasing solar flux and colder planetary body temperatures at large AU. This paper presents the challenges and opportunities of passive cryogenic cooling versus active cooling for solar powered planetary missions.

Rodriguez, Jose I.↗
NASA NTRS1992

Unsteady pressure loads in a generic high speed engine model

Unsteady pressure loads were measured along the top interior wall of a generic high-speed engine (GHSE) model undergoing performance tests in the combustion-Heated Scramjet Test Facility at the Langley Research Center. Flow to the model inlet was simulated at 72000 ft and a flight Mach number of 4. The inlet Mach number was 3.5 with a total temperature and pressure of 1640 R and 92 psia. The unsteady pressure loads were measured with 5 piezoresistive gages, recessed into the wall 4 to 12 gage diameters to reduce incident heat flux to the diaphragms, and distributed from the inlet to the combustor. Contributors to the unsteady pressure loads included boundary layer turbulence, combustion noise, and transients generated by unstart loads. Typical turbulent boundary layer rms pressures in the inlet ranged from 133 dB in the inlet to 181 dB in the combustor over the frequency range from 0 to 5 kHz. Downstream of the inlet exist, combustion noise was shown to dominate boundary layer turbulence noise at increased heat release rates. Noise levels in the isolator section increased by 15 dB when the fuel-air ratio was increased from 0.37 to 0.57 of the stoichiometric ratio. Transient pressure disturbances associated with engine unstarts were measured in the inlet and have an upstream propagation speed of about 7 ft/sec and pressure jumps of at least 3 psia.

Parrott, Tony L.↗
NASA NTRS2014

Cryogenic Boil-Off Reduction System Testing

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration due to the high specific impulse that can be achieved using engines suitable for moving 10's to 100's of metric tons of payload mass to destinations outside of low earth orbit. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several days. The losses can be greatly reduced by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and by the integration of self-supporting multi-layer insulation. The active thermal control technology under development is the integration of the reverse turbo- Brayton cycle cryocooler to the propellant tank through a distributed cooling network of tubes coupled to a shield in the tank insulation and to the tank wall itself. Also, the self-supporting insulation technology was utilized under the shield to obtain needed tank applied LH2 performance. These elements were recently tested at NASA Glenn Research Center in a series of three tests, two that reduced LH2 boil-off and one to eliminate LO2 boil-off. This test series was conducted in a vacuum chamber that replicated the vacuum of space and the temperatures of low Earth orbit. The test results show that LH2 boil-off was reduced 60% by the cryocooler system operating at 90K and that robust LO2 zero boil-off storage, including full tank pressure control was achieved.

Propellant Storage↗
NASA NTRS2012

Large Eddy/Reynolds-Averaged Navier-Stokes Simulations of CUBRC Base Heating Experiments

ven with great advances in computational techniques and computing power during recent decades, the modeling of unsteady separated flows, such as those encountered in the wake of a re-entry vehicle, continues to be one of the most challenging problems in CFD. Of most interest to the aerothermodynamics community is accurately predicting transient heating loads on the base of a blunt body, which would result in reduced uncertainties and safety margins when designing a re-entry vehicle. However, the prediction of heat transfer can vary widely depending on the turbulence model employed. Therefore, selecting a turbulence model which realistically captures as much of the flow physics as possible will result in improved results. Reynolds Averaged Navier Stokes (RANS) models have become increasingly popular due to their good performance with attached flows, and the relatively quick turnaround time to obtain results. However, RANS methods cannot accurately simulate unsteady separated wake flows, and running direct numerical simulation (DNS) on such complex flows is currently too computationally expensive. Large Eddy Simulation (LES) techniques allow for the computation of the large eddies, which contain most of the Reynolds stress, while modeling the smaller (subgrid) eddies. This results in models which are more computationally expensive than RANS methods, but not as prohibitive as DNS. By complimenting an LES approach with a RANS model, a hybrid LES/RANS method resolves the larger turbulent scales away from surfaces with LES, and switches to a RANS model inside boundary layers. As pointed out by Bertin et al., this type of hybrid approach has shown a lot of promise for predicting turbulent flows, but work is needed to verify that these models work well in hypersonic flows. The very limited amounts of flight and experimental data available presents an additional challenge for researchers. Recently, a joint study by NASA and CUBRC has focused on collecting heat transfer data on the backshell of a scaled model of the Orion Multi-Purpose Crew Vehicle (MPCV). Heat augmentation effects due to the presence of cavities and RCS jet firings were also investigated. The high quality data produced by this effort presents a new set of data which can be used to assess the performance of CFD methods. In this work, a hybrid LES/RANS model developed at North Carolina State University (NCSU) is used to simulate several runs from these experiments, and evaluate the performance of high fidelity methods as compared to more typical RANS models. .

Salazar, Giovanni↗
NASA NTRS1982

Power Controller

The device called the Power Factor Controller (PFC) offers exceptional energy conservation potential by virtue of its ability to sense shifts in the relationship between voltage and current flow, and to match them with the motor's need. Originating from the solar heating/cooling program, the PFC senses a light load, it cuts the voltage level to the minimum needed which in turn reduces current flow and heat loss. Laboratory tests showed that the PFC could reduce power used by six to eight percent under normal motor loads, and as much as 65 percent when the motor was idling. Over 150 companies have been granted NASA licenses for commercial use of this technology. One system that utilizes this technology is the Vectrol Energy System, (VES) produced by Vectrol, Inc. a subsidiary of Westinghouse. The VES is being used at Woodward & Lothrop, on their escalators. Energy use is regulated according to how many people are on the escalator at any time. It is estimated that the energy savings are between 30 to 40 percent.

Source record↗
NASA NTRS2015

Adaptable Single Active Loop Thermal Control System (TCS) for Future Space Missions

This presentation will examine the development of a thermal control system (TCS) for future space missions utilizing a single active cooling loop. The system architecture enables the TCS to be reconfigured during the various mission phases to respond, not only to varying heat load, but to heat rejection temperature as well. The system will consist of an accumulator, pump, cold plates (evaporators), condenser radiator, and compressor, in addition to control, bypass and throttling valves. For cold environments, the heat will be rejected by radiation, during which the compressor will be bypassed, reducing the system to a simple pumped loop that, depending on heat load, can operate in either a single-phase liquid mode or two-phase mode. For warmer environments, the pump will be bypassed, enabling the TCS to operate as a heat pump. This presentation will focus on recent findings concerning two-phase flow regimes, pressure drop, and heat transfer coefficient trends in the cabin and avionics micro-channel heat exchangers when using the heat pump mode. Also discussed will be practical implications of using micro-channel evaporators for the heat pump.

two phase heat transfer↗
NASA NTRS1990

Ambient Pressure Test Rig Developed for Testing Oil-Free Bearings in Alternate Gases and Variable Pressures

The Oil-Free Turbomachinery research team at the NASA Glenn Research Center is conducting research to develop turbomachinery systems that utilize high-speed, high temperature foil (air) bearings that do not require an oil lubrication system. Such systems combine the most advanced foil bearings from industry with NASA-developed hightemperature solid-lubricant technology. New applications are being pursued, such as Oil- Free turbochargers, auxiliary power units, and turbine propulsion systems for aircraft. An Oil-Free business jet engine, for example, would be simpler, lighter, more reliable, and less costly to purchase and maintain than current engines. Another application is NASA's Prometheus mission, where gas bearings will be required for the closed-cycle turbine based power-conversion system of a nuclear power generator for deep space. To support these applications, Glenn's Oil-Free Turbomachinery research team developed the Ambient Pressure Test Rig. Using this facility, researchers can load and heat a bearing and evaluate its performance with reduced air pressure to simulate high altitude conditions. For the nuclear application, the test chamber can be purged with gases such as helium to study foil gas bearing operation in working fluids other than air.

Bauman, Steven W.↗
NASA NTRS2011

Modeling Geometric Arrangements of TiO2-Based Catalyst Substrates and Isotropic Light Sources to Enhance the Efficiency of a Photocatalystic Oxidation (PCO) Reactor

The closed confined environments of the ISS, as well as in future spacecraft for exploration beyond LEO, provide many challenges to crew health. One such challenge is the availability of a robust, energy efficient, and re-generable air revitalization system that controls trace volatile organic contaminants (VOCs) to levels below a specified spacecraft maximum allowable concentration (SMAC). Photocatalytic oxidation (PCO), which is capable of mineralizing VOCs at room temperature and of accommodating a high volumetric flow, is being evaluated as an alternative trace contaminant control technology. In an architecture of a combined air and water management system, placing a PCO unit before a condensing heat exchanger for humidity control will greatly reduce the organic load into the humidity condensate loop ofthe water processing assembly (WPA) thereby enhancing the life cycle economics ofthe WPA. This targeted application dictates a single pass efficiency of greater than 90% for polar VOCs. ~ Although this target was met in laboratory bench-scaled reactors, no commercial or SBIR-developed prototype PCO units examined to date have achieved this goal. Furthermore, the formation of partial oxidation products (e.g., acetaldehyde) was not eliminated. It is known that single pass efficiency and partial oxidation are strongly dependent upon the contact time and catalyst illumination, hence the requirement for an efficient reactor design. The objective of this study is to maximize the apparent contact time and illuminated catalyst surface area at a given reactor volume and volumetric flow. In this study, a Ti02-based photocatalyst is assumed to be immobilized on porous substrate panels and illumination derived from linear isotropic light sources. Mathematical modeling using computational fluid dynamics (CFD) analyses were performed to investigate the effect of: 1) the geometry and configuration of catalyst-coated substrate panels, 2) porosity of the supporting substrate, and 3) varying the light source and spacing on contact time and illuminated catalyst area.

Richards, Jeffrey T.↗
NASA NTRS1973

Heat transfer coefficients of dilute flowing gas-solids suspensions

Heat transfer coefficients of air-glass, argon-glass, and argon-aluminum suspensions were measured in horizontal and vertical tubes. The glass, 21.6 and 36.0 micron diameter particles, was suspended at gas Reynolds numbers between 11,000 and 21,000 and loading ratios between 0 and 2.5. The presence of particles generally reduced the heat transfer coefficient. The circulation of aluminum powder in the 0.870 inch diameter closed loop system produced tenacious deposits on protuberances into the stream. In the vertical test section, the Nusselt number reduction was attributed to viscous sublayer thickening; in the horizontal test section to particle deposition.

Kane, R. S.↗
NASA NTRS1978

Analytical evaluation of the impact of broad specification fuels on high bypass turbofan engine combustors

The impact of the use of broad specification fuels on the design, performance durability, emissions and operational characteristics of combustors for commercial aircraft gas turbine engines was assessed. Single stage, vorbix and lean premixed prevaporized combustors, in the JT9D and an advanced energy efficient engine cycle were evaluated when operating on Jet A and ERBS (Experimental Referee Broad Specification) fuels. Design modifications, based on criteria evolved from a literature survey, were introduced and their effectiveness at offsetting projected deficiencies resulting from the use of ERBS was estimated. The results indicate that the use of a broad specification fuel such as ERBS, will necessitate significant technology improvements and redesign if deteriorated performance, durability and emissions are to be avoided. Higher radiant heat loads are projected to seriously compromise liner life while the reduced thermal stability of ERBS will require revisions to the engine-airframe fuel system to reduce the thermal stress on the fuel. Smoke and emissions output are projected to increase with the use of broad specification fuels. While the basic geometry of the single stage and vorbix combustors are compatible with the use of ERBS, extensive redesign of the front end of the lean premixed prevaporized burner will be required to achieve satisfactory operation and optimum emissions.

Lohmann, R. P.↗
NASA NTRS1979

The effects of air leaks on solar air heating systems

This paper presents the results of an investigation to determine the effects of leakages in collector and duct work on the system performance of a typical single-family residence solar air heating system. Positive (leakage out) and negative (leakage in) pressure systems were examined. Collector and duct leakage rates were varied from 10 to 30 percent of the system flow rate. Within the range of leakage rates investigated, solar contribution to heated space and domestic hot water loads was found to be reduced up to 30 percent from the no-leak system contribution with duct leakage equally divided between supply and return duct; with supply duct leakage greater than return leakage a reduction of up to 35 percent was noted. The negative pressure system exhibited a reduction in solar contribution somewhat larger than the positive pressure system for the same leakage rates.

Elkin, R.↗
NASA NTRS2012

Model Development and Experimental Validation of the Fusible Heat Sink Design for Exploration Vehicles

The Fusible Heat Sink is a novel vehicle heat rejection technology which combines a flow through radiator with a phase change material. The combined technologies create a multi-function device able to shield crew members against Solar Particle Events (SPE), reduce radiator extent by permitting sizing to the average vehicle heat load rather than to the peak vehicle heat load, and to substantially absorb heat load excursions from the average while constantly maintaining thermal control system setpoints. This multi-function technology provides great flexibility for mission planning, making it possible to operate a vehicle in hot or cold environments and under high or low heat load conditions for extended periods of time. This paper describes the model development and experimental validation of the Fusible Heat Sink technology. The model developed was intended to meet the radiation and heat rejection requirements of a nominal MMSEV mission. Development parameters and results, including sizing and model performance will be discussed. From this flight-sized model, a scaled test-article design was modeled, designed, and fabricated for experimental validation of the technology at Johnson Space Center thermal vacuum chamber facilities. Testing showed performance comparable to the model at nominal loads and the capability to maintain heat loads substantially greater than nominal for extended periods of time.

Cognata, Thomas J.↗
NASA NTRS2013

Model Development and Experimental Validation of the Fusible Heat Sink Design for Exploration Vehicles

The Fusible Heat Sink is a novel vehicle heat rejection technology which combines a flow through radiator with a phase change material. The combined technologies create a multi-function device able to shield crew members against Solar Particle Events (SPE), reduce radiator extent by permitting sizing to the average vehicle heat load rather than to the peak vehicle heat load, and to substantially absorb heat load excursions from the average while constantly maintaining thermal control system setpoints. This multi-function technology provides great flexibility for mission planning, making it possible to operate a vehicle in hot or cold environments and under high or low heat load conditions for extended periods of time. This paper describes the modeling and experimental validation of the Fusible Heat Sink technology. The model developed was intended to meet the radiation and heat rejection requirements of a nominal MMSEV mission. Development parameters and results, including sizing and model performance will be discussed. From this flight-sized model, a scaled test-article design was modeled, designed, and fabricated for experimental validation of the technology at Johnson Space Center thermal vacuum chamber facilities. Testing showed performance comparable to the model at nominal loads and the capability to maintain heat loads substantially greater than nominal for extended periods of time.

Cognata, Thomas J.↗
NASA NTRS2016

Thermal Vacuum Testing of a Helium Loop Heat Pipe for Large Area Cryocooling

Future NASA space telescopes and exploration missions require cryocooling of large areas such as optics, detector arrays, and cryogenic propellant tanks. One device that can potentially be used to provide closed-loop cryocooling is the cryogenic loop heat pipe (CLHP). A CLHP has many advantages over other devices in terms of reduced mass, reduced vibration, high reliability, and long life. A helium CLHP has been tested extensively in a thermal vacuum chamber using a cryocooler as the heat sink to characterize its transient and steady performance and to verify its ability to cool large areas or components in the 3 degrees Kelvin temperature range. The helium CLHP thermal performance test included cool-down from the ambient temperature, startup, capillary limit, heat removal capability, rapid power changes, and long duration steady state operation. The helium CLHP demonstrated robust operation under steady state and transient conditions. The loop could be cooled from the ambient temperature to subcritical temperatures very effectively, and could start successfully by simply applying power to both the capillary pump and the evaporator plate without pre-conditioning. It could adapt to a rapid heat load change and quickly reach a new steady state. Heat removal between 10 megawatts and 140 megawatts was demonstrated, yielding a power turn down ratio of 14. When the CLHP capillary limit was exceeded, the loop could resume its normal function by reducing the power to the capillary pump. Steady state operations up to 17 hours at several heat loads were demonstrated. The ability of the helium CLHP to cool large areas was therefore successfully verified.

Thermal Control Systems↗
NASA NTRS2015

Testing of a Helium Loop Heat Pipe for Large Area Cryocooling

Future NASA space telescopes and exploration missions require cryocooling of large areas such as optics, detector arrays, and cryogenic propellant tanks. One device that can potentially be used to provide closed-loop cryocooling is the cryogenic loop heat pipe (CLHP). A CLHP has many advantages over other devices in terms of reduced mass, reduced vibration, high reliability, and long life. A helium CLHP has been tested extensively in a thermal vacuum chamber using a cryocooler as the heat sink to characterize its transient and steady performance and verify its ability to cool large areas or components in the 3K temperature range. A copper plate with attached electrical heters was used to simulate the heat source, and heat was collected by the CLHP evaporator and transferred to the cryocooler for ultimate heat rejection. The helium CLHP thermal performance test included cool-down from the ambient temperature, startup, capillary limit, heat removal capability, rapid power changes, and long duration steady state operation. The helium CLHP demonstrated robust operation under steady state and transient conditions. The loop could be cooled from the ambient temperature to subcritical temperatures very effectively, and could start successfully without pre-conditioning by simply applying power to both the capillary pump and the evaporator plate. It could adapt to rapid changes in the heat load, and reach a new steady state very quickly. Heat removal between 10mW and 140mW was demonstrated, yielding a power turn down ratio of 14. When the CLHP capillary limit was exceeded, the loop could resume its normal function by reducing the power to the capillary pump. Steady state operations up to 17 hours at several heat loads were demonstrated. The ability of the helium CLHP to cool large areas was therefore successfully verified.

Crycooling↗