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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.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
DOE OSTI2023

Strongly Heated Turbulent Flow in a Channel with Pin Fins

Large-eddy simulations (LES) were performed to study the turbulent flow in a channel of height H with a staggered array of pin fins with diameter D = H/2 as a function of heating loads that are relevant to the cooling of turbine blades and vanes. The following three heating loads were investigated—wall-to-coolant temperatures of Tw/Tc = 1.01, 2.0, and 4.0 - where the Reynolds number at the channel inlet was 10,000 and the back pressure at the channel outlet was 1 bar. For the LES, two different subgrid-scale models—the dynamic kinetic energy model (DKEM) and the wall-adapting local eddy-viscosity model (WALE)—were examined and compared. This study was validated by comparing with data from direct numerical simulation and experimental measurements. The results obtained show high heating loads to create wall jets next to all heated surfaces that significantly alter the structure of the turbulent flow. Results generated on effects of heat loads on the mean and fluctuating components of velocity and temperature, turbulent kinetic energy, the anisotropy of the Reynolds stresses, and velocity-temperature correlations can be used to improve existing RANS models.

42 ENGINEERING↗
NASA NTRS2008

Review of End-of-Life Thermal Control Coating Performance

White thermal control coatings capable of long term performance are needed for Fission Surface Power (FSP) where heat from a nuclear reactor placed on the surface of the Moon must be rejected to the environment. The threats to thermal control coating durability on the lunar surface are electrons, protons, and ultraviolet radiation. The anticipated damage to the coating is a gradual darkening over time. The increase in solar absorptance would, in essence, add a cyclic heat load to the radiator. The greater the darkening, the greater the added heat load. The cyclic heat load could ultimately impart a cyclic influence on FSP system performance. No significant change in emittance is anticipated. Optical properties degradation data were found in the open literature for the Z-93 series of thermal control paints. Additional optical properties degradation data were found from the Lunar Orbiter V mission, the Optical Properties Monitor, and the Materials International Space Station Experiment. Anticipated end-of-life thermal control coating performance for a FSP installation is postulated. With the FSP installation located away from landing and launching areas, and out of line-of-sight, lunar dust from human activity may not be a threat. The benefits of investing in next generation thermal control paint chemistry are explored.

Jaworske, Donald A.↗
DOE OSTI2025

Real time computations of cryogenic He properties

The Fermilab PIP-II (proton improvement plan - II) project is being constructed at Fermilab to deliver $800\,MeV$ protons of $>1\,MW$ beam power to replace the present LINAC and provide protons to the remainder of the existing accelerator complex. The new LINAC consists of a warm front end, 23 superconducting RF cryomodules, and a beam transfer line to the existing complex. The cryomodules (CMs) are to be tested at Fermilab's CryoModule Test Facility (CMTF).An important measurement in cryogenic testing is the heat load of each CM. Traditionally, at Fermilab, these measurements were made collecting archived data offline and analyzing it. The new control system for PIP-II is being developed with the EPICS (Experimental Physics and Industrial Control System) framework, which allows us to compute the heat load in real time using the HePak library.We are exploring other $He$ properties, such as flow, where flow meters are not available, which can also be calculated in real time and fed back to the cryogenics engineers.This paper details the real time heat load calculation and $He$ flow software developed for CM testing at CMTF, as well as the first results from the prototype HB650 CM. Future plans for 2-phase $LHe$ flow will also be outlined.

Hanlet, Pierrick [Fermilab]↗
NASA NTRS2008

Spacecraft Radiator Freeze Protection Using a Regenerative Heat Exchanger with Bypass Setpoint Temperature Control

Spacecraft radiators are sized for their maximum heat load in their warmest thermal environment, but must operate at reduced heat loads and in colder environments. For systems where the radiator environment can be colder than the working fluid freezing temperature, radiator freezing becomes an issue. Radiator freezing has not been a major issue for the Space Shuttle and the International Space Station (ISS) active thermal control systems (ATCSs) because they operate in environments that are warm relative to the freezing point of their external coolants (Freon-21 and ammonia, respectively). For a vehicle that lands at the Lunar South Pole, the design thermal environment is 215K, but the radiator working fluid must also be kept from freezing during the 0 K sink of transit. A radiator bypass flow control design such as those used on the Space Shuttle and ISS requires more than 30% of the design heat load to avoid radiator freezing during transit - even with a very low freezing point working fluid. By changing the traditional ATCS architecture to include a regenerating heat exchanger inboard of the radiator and by using a regenerator bypass flow control valve to maintain system setpoint, the required minimum heat load can be reduced by more than half. This gives the spacecraft much more flexibility in design and operation. The present work describes the regenerator bypass ATCS setpoint control methodology. It includes analytical results comparing the performance of this system to the traditional radiator bypass system. Finally, a summary of the advantages of the regenerator bypass system are presented.

Ungar, Eugene K.↗
DOE OSTI2017

PIP-II Cryogenic System and the evolution of Superfluid Helium Cryogenic Plant Specifications

The PIP-II cryogenic system consists of a Superfluid Helium Cryogenic Plant (SHCP) and a Cryogenic Distribution System (CDS) connecting the SHCP to the Superconducting (SC) Linac consisting of 25 cryomodules. The dynamic heat load of the SC cavities for continuous wave (CW) as well as pulsed mode of operation has been listed out. The static heat loads of the cavities along with the CDS have also been discussed. Simulation study has been carried out to compute the supercritical helium (SHe) flow requirements for each cryomodule. Comparison between the flow requirements of the cryomodules for the CW and pulsed modes of operation have also been made. From the total computed heat load and pressure drop values in the CDS, the basic specifications for the SHCP, required for cooling the SC Linac, have evolved.

43 PARTICLE ACCELERATORS↗
NASA NTRS2017

Testing Seam Concepts for Advanced Multilayer Insulation

Multilayer insulation (MLI) is considered the state of the art insulation for cryogenic propellant tanks in the space environment. MLI traditionally consists of multiple layers of metalized films separated by low conductivity spacers. In order to better understand some of the details within MLI design and construction, GRC has been investigating the heat loads caused by multiple types of seams. To date testing has been completed with 20 layer and 50 layer blankets. Although a truly seamless blanket is not practical, a blanket lay-up where each individual layer was overlapped and tapped together was used as a baseline for the other seams tests. Other seams concepts tested included: an overlap where the complete blanket was overlapped on top of itself; a butt joint were the blankets were just trimmed and butted up against each other, and a staggered butt joint where the seam in the out layers is offset from the seam in the inner layers. Measured performance is based on a preliminary analysis of rod calibration tests conducted prior to the start of seams testing. Baseline performance for the 50 layer blanket showed a measured heat load of 0.46 Watts with a degradation to about 0.47 Watts in the seamed blankets. Baseline performance for the 20 layer blanket showed a measured heat load of 0.57 Watts. Heat loads for the seamed tests are still begin analyzed. So far analysis work has suggested the need for corrections due to heat loads from both the heater leads and the instrumentation wires. A careful re-examination of the calibration test results with these factors accounted for is also underway. This presentation will discuss the theory of seams in MLI, our test results to date, and the uncertainties in our measurements.

multilayer insulation↗
NASA NTRS1986

Aerothermodynamics of transatmospheric vehicles

A transatmospheric vehicle (TAV) using airbreathing propulsion requires a long acceleration period within the denser part of the atmosphere to reach orbital speed. The long flight time, coupled with the need for a low-drag configuration, results in severe heating of parts of the vehicle. The ascent peak stagnation point and wing leading edge equilibrium wall temperatures are about 3500 K and 2500 K, respectively, likely requiring some form of mass addition cooling. The corresponding temperatures during entry are 1000 K lower. The vehicle windward centerline temperatures are more moderate, with values peaking around 1300 K during both ascent and entry. Therefore, radiative cooling should be effective over large areas of the vehicle. The windward centerline heat loads during entry are comparable to those for low acceleration ascent trajectories. However, ascent heat loads for the stagnation point and the wing leading edge are about three times higher than those during entry. For comparison, the entry heat load for the TAV's stagnation point is about three times higher than the value for Shuttle. Therefore the ascent heat load at the TAV's stagnation point exceeds the Shuttle's entry value by an order of magnitude.

Tauber, M. E.↗
DOE OSTI2015

Long pulse acceleration of MeV class high power density negative H{sup −} ion beam for ITER

R and D of high power density negative ion beam acceleration has been carried out at MeV test facility in JAEA to realize ITER neutral beam accelerator. The main target is H{sup −} ion beam acceleration up to 1 MeV with 200 A/m{sup 2} for 60 s whose pulse length is the present facility limit. For long pulse acceleration at high power density, new extraction grid (EXG) has been developed with high cooling capability, which electron suppression magnet is placed under cooling channel similar to ITER. In addition, aperture size of electron suppression grid (ESG) is enlarged from 14 mm to 16 mm to reduce direct interception on the ESG and emission of secondary electron which leads to high heat load on the upstream acceleration grid. By enlarging ESG aperture, beam current increased 10 % at high current beam and total acceleration grid heat load reduced from 13 % to 10 % of input power at long pulse beam. In addition, heat load by back stream positive ion into the EXG is measured for the first time and is estimated as 0.3 % of beam power, while heat load by back stream ion into the source chamber is estimated as 3.5 ~ 4.0 % of beam power. Beam acceleration up to 60 s which is the facility limit, has achieved at 683 keV, 100 A/m{sup 2} of negative ion beam, whose energy density increases two orders of magnitude since 2011.

43 PARTICLE ACCELERATORS↗
NASA NTRS1988

Cryogenic systems for the large deployable reflector

There are five technologies which may have application for Large Deployable Reflector (LDR), one passive and four active. In order of maturity, they are passive stored cryogen systems, and mechanical, sorption, magnetic, and pulse-tube refrigerators. In addition, deep space radiators will be required to reject the heat of the active systems, and may be useful as auxiliary coolers for the stored cryogen systems. Hybrid combinations of these technologies may well be more efficient than any one alone, and extensive system studies will be required to determine the best trade-offs. Stored cryogen systems were flown on a number of missions. The systems are capable of meeting the temperature requirements of LDR. The size and weight of stored cryogen systems are proportional to heat load and, as a result, are applicable only if the low-temperature heat load can be kept small. Systems using chemisorption and physical adsorption for compressors and pumps have received considerable attention in the past few years. Systems based on adiabatic demagnetization of paramagnetic salts were used for refrigeration for many years. Pulse-tube refrigerators were recently proposed which show relatively high efficiency for temperatures in the 60 to 80 K range. The instrument heat loads and operating temperatures are critical to the selection and design of the cryogenic system. Every effort should be made to minimize heat loads, raise operating temperatures, and to define these precisely. No one technology is now ready for application to LDR. Substantial development efforts are underway in all of the technologies and should be monitored and advocated. Magnetic and pulse-tube refrigerators have high potential.

Mason, Peter V.↗
NASA NTRS2009

Aerothermodynamic Design of the Mars Science Laboratory Backshell and Parachute Cone

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule backshell and parachute cone. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux) design entry trajectories from a 2009 launch. Transient interference effects from reaction control system thruster plumes were included in the design environments when necessary. The limiting backshell design heating conditions of 6.3 W/sq cm for heat flux and 377 J/sq cm for total heat load are not influenced by thruster firings. Similarly, the thrusters do not affect the parachute cover lid design environments (13 W/sq cm and 499 J/sq cm). If thruster jet firings occur near peak dynamic pressure, they will augment the design environments at the interface between the backshell and parachute cone (7 W/sq cm and 174 J/sq cm). Localized heat fluxes are higher near the thruster fairing during jet firings, but these areas did not require additional thermal protection material. Finally, heating bump factors were developed for antenna radomes on the parachute cone

Edquist, Karl T.↗
NASA NTRSDate not supplied

Results of Use of Heat Flux Sensors on Liquid Hydrogen Tanks

Heat flux sensors were used to characterize the performance of insulation on the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER), a large-scale test article designed to simulate upper stage cryogenic propellant tank thermal performance in simulated space environments. Usually, the insulation heat loads are derived from calculations removing all other heat sources and attributing the residual heat load to the insulation system. Testing for SHIIVER included the tank being just insulated with spray-on-foam insulation as well as covering the domes with multilayer insulation while leaving the barrel section insulated with spray-on-foam. Heat flux sensors were located at multiple locations on both domes as well as on the barrel section of the tank. Results from the SHIIVER testing using the heat flux sensors are compared to other calculated heat inputs for both liquid nitrogen and liquid hydrogen testing as a function of tank fill level. Further investigation into the transient nature of the SHIIVER testing including the heat flux sensors provided insight into heat flow patterns that may not have been otherwise seen using temperature sensors and calculated insulation heat loads. While the demonstrated uncertainties in the absolute values in the heat flux sensors are high, the values and the trends match well with other calculation methods. The results of SHIIVER allow for the use of heat flux sensors for measurements of insulation performance and dynamic system thermal response for future applications.

Cryogenic Fluid Management↗
NASA NTRSDate not supplied

Results of Use of Heat Flux Sensors on Liquid Hydrogen Tanks

ABSTRACT: Heat flux sensors were used to characterize the performance of insulation on the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER), a large-scale test article designed to simulate upper stage cryogenic propellant tank thermal performance in simulated space environments. Usually, the insulation heat loads are derived from calculations removing all other heat sources and attributing the residual heat load to the insulation system. Testing for SHIIVER included the tank being just insulated with spray-on-foam insulation as well as covering the domes with multilayer insulation while leaving the barrel section insulated with spray-on-foam. Heat flux sensors were located at multiple locations on both domes as well as on the barrel section of the tank. Results from the SHIIVER testing using the heat flux sensors are compared to other calculated heat inputs for both liquid nitrogen and liquid hydrogen testing as a function of tank fill level. Further investigation into the transient nature of the SHIIVER testing including the heat flux sensors provided insight into heat flow patterns that may not have been otherwise seen using temperature sensors and calculated insulation heat loads. While the demonstrated uncertainties in the absolute values in the heat flux sensors are high, the values and the trends match well with other calculation methods. The results of SHIIVER allow for the use of heat flux sensors for measurements of insulation performance and dynamic system thermal response for future applications.

Cryogenic Fluid Management↗
DOE OSTI2021

Cost-Optimized Cold Climate Heat Pump Development and Field Test

Cold climate heat pumps (CCHPs) expand the heat pump market to climates where heating demand is dominant. They can achieve more than 70% energy savings compared with electric resistance heating and operate at lower cost than using tank-stored propane to fuel a furnace. A high-efficiency heat pump with a heating seasonal performance factor (HSPF)—as defined by the Air-Conditioning, Heating, and Refrigeration Institute (AHRI 2008)—greater than 10.0 would be more efficient than gas heating in terms of source energy. However, developing a cost-effective CCHP involves some challenges. A typical single-speed, air-source heat pump (ASHP) with an HSPF of 7.7 Btu/Wh does not work well under cold outdoor temperature conditions typical of cold climate locations for four major reasons: 1. Discharge temperature is too high—The low suction pressure and high compression pressure ratio at low ambient temperatures causes high compressor discharge temperatures in excess of the maximum limit for many of the compressors on the market. Furthermore, system charge of a heat pump is usually optimized in the cooling mode, which leads to overcharge conditions in the heating mode, further increasing the discharge temperature. 2. Heating capacity is insufficient if sized to meet the building design cooling load— Heating capacity of a single-speed heat pump decreases with ambient temperature. The heating capacity at -13°F (-25°C) typically decreases to 20%–40% of the rated heating capacity at 47°F (8.3°C) (~equivalent to the rated cooling capacity at 95°F [35°C]). Therefore, a single-speed heat pump, typically sized to match the building design cooling load, cannot provide adequate heating capacity to match the building heating load at low ambient temperatures. The capacity deficit is filled by inefficient resistance heat, thus lowering the system efficiency and significantly increasing power demand. 3. Cyclic loss is significant if sized to meet the building design heating load—If a single-speed heat pump is sized to meet the heating load, it will be significantly oversized relative to the cooling load in many cold climates. This will cause excessive on/off cyclic loss during the cooling and heating operations at moderately low ambient temperatures. Thus, capacity modulation capability (e.g., using a variable-speed or multi-stage compressor) is necessary for a CCHP, which uses its full capacity to meet the peak heating load and partial capacity to meet the cooling and part-load heating loads. 4. Coefficient of performance (COP) is low—Heating COP degrades significantly at low ambient temperatures owing to the large temperature difference between the heat source and sink. A target CCHP should be sized to meet the building design heating load while minimizing the cyclic loss for the cooling and heating operations at moderate ambient temperatures.

32 ENERGY CONSERVATION, CONSUMPTION, AND UTILIZATI↗
NASA NTRS1978

Jet pump assisted arterial heat pipe

This paper discusses the concept of an arterial heat pipe with a capillary driven jet pump. The jet pump generates a suction which pumps vapor and noncondensible gas from the artery. The suction also forces liquid into the artery and maintains it in a primed condition. A theoretical model was developed which predicts the existence of two stable ranges. Up to a certain tilt the artery will prime by itself once a heat load is applied to the heat pipe. At higher tilts, the jet pump can maintain the artery in a primed condition but self-priming is not possible. A prototype heat pipe was tested which self-primed up to a tilt of 1.9 cm, with a heat load of 500 watts. The heat pipe continued to prime reliably when operated as a VCHP, i.e., after a large amount of noncondensible gas was introduced.

Bienert, W. B.↗
NASA NTRS1980

Solar heating system at Security State Bank, Starkville, Mississippi

The 312 square feet of Solaron flat plate air collectors provide for 788 square feet of space heating, an estimated 55 percent of the heating load. Solar heated air is distributed to the 96 cubic foot steel cylinder, which contains two inch diameter rocks. An air handler unit moves the air over the collector and into the steel cylinder. Four motorized dampers and two gravity dampers are also part of the system. A Solaron controller which has sensors located at the collectors, rock storage, and at the return air, automatically controls the system. Auxiliary heating energy is provided by electric resistance duct heaters.

Source record↗
NASA NTRS1985

Temperature control of silica mirrors for cryogenically-cooled telescopes in space

The SIRTF (Space Infrared Telescope Facility) optical subsystem, including the primary mirror, has been modeled using lumped-parameter techniques and a thermal and cryogenic transient analysis program developed previously. Since silica mirrors have been shown to have significantly less distortion due to cryogenic cooldown, a silica primary mirror of approximately 1 m diameter is being considered. Thermal responses to variations in aperture heat load and heat sink temperature have been explored. Calculations have been made of system noise components from primary mirror spatial and temporal temperature variations. Background-limited infrared observations can be made at wavelengths out to 200 micrometers. Mirror thermal performance with and without copper wires for cooling is also considered.

Melugin, R. K.↗
NASA NTRS2007

A Comparison of Two Skip Entry Guidance Algorithms

The Orion capsule vehicle will have a Lift-to-Drag ratio (L/D) of 0.3-0.35. For an Apollo-like direct entry into the Earth's atmosphere from a lunar return trajectory, this L/D will give the vehicle a maximum range of about 2500 nm and a maximum crossrange of 216 nm. In order to y longer ranges, the vehicle lift must be used to loft the trajectory such that the aerodynamic forces are decreased. A Skip-Trajectory results if the vehicle leaves the sensible atmosphere and a second entry occurs downrange of the atmospheric exit point. The Orion capsule is required to have landing site access (either on land or in water) inside the Continental United States (CONUS) for lunar returns anytime during the lunar month. This requirement means the vehicle must be capable of flying ranges of at least 5500 nm. For the L/D of the vehicle, this is only possible with the use of a guided Skip-Trajectory. A skip entry guidance algorithm is necessary to achieve this requirement. Two skip entry guidance algorithms have been developed: the Numerical Skip Entry Guidance (NSEG) algorithm was developed at NASA/JSC and PredGuid was developed at Draper Laboratory. A comparison of these two algorithms will be presented in this paper. Each algorithm has been implemented in a high-fidelity, 6 degree-of-freedom simulation called the Advanced NASA Technology Architecture for Exploration Studies (ANTARES). NASA and Draper engineers have completed several monte carlo analyses in order to compare the performance of each algorithm in various stress states. Each algorithm has been tested for entry-to-target ranges to include direct entries and skip entries of varying length. Dispersions have been included on the initial entry interface state, vehicle mass properties, vehicle aerodynamics, atmosphere, and Reaction Control System (RCS). Performance criteria include miss distance to the target, RCS fuel usage, maximum g-loads and heat rates for the first and second entry, total heat load, and control system saturation. The comparison of the performance criteria has led to a down select and guidance merger that will take the best ideas from each algorithm to create one skip entry guidance algorithm for the Orion vehicle.

Rea, Jeremy R.↗