Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “view factor”

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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Pupil geometry and pupil re-imaging in telescope arrays

This paper considers the issues of lateral and longitudinal pupil geometry in ground-based telescope arrays, such as IOTA. In particular, it is considered whether or not pupil re-imaging is required before beam combination. By considering the paths of rays through the system, an expression is derived for the optical path errors in the combined wavefront as a function of array dimensions, telescope magnification factor, viewing angle, and field-of-view. By examining this expression for the two cases of pupil-plane and image-plane combination, operational limits can be found for any array. As a particular example, it is shown that for IOTA no pupil re-imaging optics will be needed.

Traub, Wesley A.↗
NASA NTRS1994

Reference manual for the Thermal Analyst's Help Desk Expert System

This document provides technical information and programming guidance for the maintenance and future development of the Thermal Analyst's Help Desk. Help Desk is an expert system that operates within the EXSYSTM expert system shell, and is used to determine first approximations of thermal capacity for spacecraft and instruments. The five analyses supported in Help Desk are: (1) surface area required for a radiating surface, (2) equilibrium temperature of a surface, (3) enclosure temperature and heat loads for a defined position in orbit, (4) enclosure temperature and heat loads over a complete orbit and, (5) selection of appropriate surface properties. The two geometries supported by Help Desk are a single flat plate and a rectangular box enclosure. The technical information includes the mathematical approach and analytical derivations used in the analyses such as: radiation heat balance, view factor calculation, and orbit determination with coordinate transformation. The programming guide for developers describes techniques for enhancement of Help Desk. Examples are provided showing the addition of new features, user interface development and enhancement, and external program interfaces.

Ormsby, Rachel A.↗
NASA NTRS1994

ASTM E 1559 method for measuring material outgassing/deposition kinetics has applications to aerospace, electronics, and semiconductor industries

The American Society for Testing and Materials has published a new standard test method for characterizing time and temperature-dependence of material outgassing kinetics and the deposition kinetics of outgassed species on surfaces at various temperatures. This new ASTM standard, E 1559(1), uses the quartz crystal microbalance (QCM) collection measurement approach. The test method was originally developed under a program sponsored by the United States Air Force Materials Laboratory (AFML) to create a standard test method for obtaining outgassing and deposition kinetics data for spacecraft materials. Standardization by ASTM recognizes that the method has applications beyond aerospace. In particular, the method will provide data of use to the electronics, semiconductor, and high vacuum industries. In ASTM E 1559 the material sample is held in vacuum in a temperature-controlled effusion cell, while its outgassing flux impinges on several QCM's which view the orifice of the effusion cell. Sample isothermal total mass loss (TML) is measured as a function of time from the mass collected on one of the QCM's which is cooled by liquid nitrogen, and the view factor from this QCM to the cell. The amount of outgassed volatile condensable material (VCM) on surfaces at higher temperatures is measured as a function of time during the isothermal outgassing test by controlling the temperatures of the remaining QCM's to selected values. The VCM on surfaces at temperatures in between those of the collector QCM's is determined at the end of the isothermal test by heating the QCM's at a controlled rate and measuring the mass loss from the end of the QCM's as a function of time and temperature. This reevaporation of the deposit collected on the QCM's is referred to as QCM thermogravimetric analysis. Isothermal outgassing and deposition rates can be determined by differentiating the isothermal TML and VCM data, respectively, while the evaporation rates of the species can be obtained as a function of temperature by differentiating the QCM thermogravimetric analysis data.

Garrett, J. W.↗
NASA NTRS1995

Numerical solution of fluid flow and heat tranfer problems with surface radiation

This paper presents a numerical scheme, based on the finite element method, to solve strongly coupled fluid flow and heat transfer problems. The surface radiation effect for gray, diffuse and isothermal surfaces is considered. A procedure for obtaining the view factors between the radiating surfaces is discussed. The overall solution strategy is verified by comparing the available results with those obtained using this approach. An analysis of a thermosyphon is undertaken and the effect of considering the surface radiation is clearly explained.

Ahuja, S.↗
NASA NTRS2000

Thermal Consideration of SWIFT XRT Radiator At-35C or Colder in Low Earth Orbit

The X-Ray Telescope (XRT) is an instrument on the National Aeronautics and Space Administration (NASA) SWIFT spacecraft. The thermoelectric cooler (TEC) for the charge coupled device (CCD) of the XRT requires a radiator temperature of -35 C or colder, and a goal of -55 C to minimize the damage by radiation. The waste heat rejected from the TEC to the radiator is in the 8 W to 20 W range. In the Phase A baseline design, the XRT radiator is mounted to the rear end of the XRT telescope tube and is very close to the bottom closeout of the spacecraft bus. The bottom closeout is multi-layer insulation (MLI) blankets. At sun angles between 90 deg and 180 deg, there is direct solar impingement on the bottom closeout. When the rolls +/- 5 deg, the XRT radiator is exposed to direct solar radiation. The radiator also has a view factor to the solar arrays. The results of thermal analysis showed that the flight temperature prediction of the radiator exceeds the temperature requirement of -35 C substantially at sun angles from 110 deg to 180 deg. A new location on the anti-sun side of the spacecraft is proposed for the radiator. It requires a heat pipe to couple the TEC and the radiator thermally. The results of thermal analysis show that the flight temperature prediction of the proposed radiator meets the temperature requirement at all sun angles.

Choi, Michael K.↗
NASA NTRS2001

An Analysis and Procedure for Determining Space Environmental Sink Temperatures With Selected Computational Results

The purpose of this report was to analyze the heat-transfer problem posed by the determination of spacecraft temperatures and to incorporate the theoretically derived relationships in the computational code TSCALC. The basis for the code was a theoretical analysis of the thermal radiative equilibrium in space, particularly in the Solar System. Beginning with the solar luminosity, the code takes into account these key variables: (1) the spacecraft-to-Sun distance expressed in astronomical units (AU), where 1 AU represents the average Sun-to-Earth distance of 149.6 million km; (2) the angle (arc degrees) at which solar radiation is incident upon a spacecraft surface (ILUMANG); (3) the spacecraft surface temperature (a radiator or photovoltaic array) in kelvin, the surface absorptivity-to-emissivity ratio alpha/epsilon with respect to the solar radiation and (alpha/epsilon)(sub 2) with respect to planetary radiation; and (4) the surface view factor to space F. Outputs from the code have been used to determine environmental temperatures in various Earth orbits. The code was also utilized as a subprogram in the design of power system radiators for deep-space probes.

Juhasz, Albert J.↗
NASA NTRS2008

Potential of a New Lunar Surface Radiator Concept for Hot Lunar Thermal Environments

The optimum radiator configuration in hot lunar thermal environments is one in which the radiator is parallel to the ground and has no view to the hot lunar surface. However, typical spacecraft configurations have limited real estate available for top-mounted radiators, resulting in a desire to use the spacecraft s vertically oriented sides. Vertically oriented, flat panel radiators will have a large view factor to the lunar surface, and thus will be subjected to significant incident lunar infrared heat. Consequently, radiator fluid temperatures will need to exceed approx.325 K (assuming standard spacecraft radiator optical properties) in order to provide positive heat rejection at lunar noon. Such temperatures are too high for crewed spacecraft applications in which a heat pump is to be avoided. A recent study of vertically oriented radiator configurations subjected to lunar noon thermal environments led to the discovery of a novel radiator concept that yielded positive heat rejection at lower fluid temperatures. This radiator configuration, called the Upright Lunar Terrain Radiator Assembly (ULTRA), has exhibited superior performance to all previously analyzed concepts in terms of heat rejection in the lunar noon thermal environment. A key benefit of the ULTRA is the absence of louvers or other moving parts and its simple geometry. Analysis of the ULTRA for a lunar extravehicular activity (EVA) portable life support system (PLSS) is shown to provide moderate heat rejection, on average, at all solar incident angles assuming an average radiator temperature of 294 K, whereas prior concepts exhibited insignificant heat rejection or heat absorption at higher incident angles. The performance of the ULTRA for a lunar lander is also discussed and compared to the performance of a vertically oriented, flat panel radiator at various lunar latitudes.

Ochoa, Dustin A.↗
NASA NTRS2010

Integrated Solar Concentrator and Shielded Radiator

A shielded radiator is integrated within a solar concentrator for applications that require protection from high ambient temperatures with little convective heat transfer. This innovation uses a reflective surface to deflect ambient thermal radiation, shielding the radiator. The interior of the shield is also reflective to provide a view factor to deep space. A key feature of the shield is the parabolic shape that focuses incoming solar radiation to a line above the radiator along the length of the trough. This keeps the solar energy from adding to the radiator load. By placing solar cells along this focal line, the concentration of solar energy reduces the number and mass of required cells. By shielding the radiator, the effective reject temperature is much lower, allowing lower radiator temperatures. This is particularly important for lower-temperature processes, like habitat heat rejection and fuel cell operations where a high radiator temperature is not feasible. Adding the solar cells in the focal line uses the concentrating effect of the shield to advantage to accomplish two processes with a single device. This shield can be a deployable, lightweight Mylar structure for compact transport.

Clark, David Larry↗
NASA NTRS2002

A Simplified, Closed-Form Method for Screening Spacecraft Orbital Heating Variations

A closed-form analytical technique has been developed to screen orbital average heating variations as a function of beta angle, altitude, surface area, and surface optical properties. Using planetary view factor equations for surfaces parallel-to and normal-to the local vertical, a cylindrical umbral shadow approximation, and a simplified albedo flux model, heating rate equations are formulated and then integrated to obtain orbital average heating. The results are compared to detailed analytical predictions using Monte Carlo integration and an assessment of error is presented.

Rickman, S. L.↗
NASA NTRS2010

Preservation of Thermal Control Specular Gold Baffle Surface on the James Webb Space Telescope (JWST) Integrated Science Instrument Module (ISIM) Electronics Compartment (IEC)

Extremely tight thermal control property degradation allowances on the vapor-deposited, gold-coated IEC baffle surface, made necessary by the cryogenic JWST Observatory operations, dictate tight contamination requirements on adjacent surfaces. Theoretical degradation in emittance with contaminant thickness was calculated. Maximum allowable source outgassing rates were calculated using worst case view factors from source to baffle surface. Tight requirements pushed the team to change the design of the adjacent surfaces to minimize the outgassing sources

MonteedeGarcia, Kristina↗
NASA NTRS2000

Thermal Analysis of ISS Service Module Active TCS

ISS Service Module mission must begin in July 2000. The verification of design thermal requirements is mostly due to thermal analysis. The thermal analysis is enough difficult problem because of large number of ISS configurations that had to be investigated and various orbital environments. Besides the ISS structure has articulating parts such as solar arrays and radiators. The presence of articulating parts greatly increases computation times and requires accurate approach to organization of calculations. The varying geometry needs us to calculate the view factors several times during the orbit, while in static geometry case we need do it only once. In this paper we consider the thermal mathematical model of SM that includes the TCS and construction thermal models and discuss the results of calculations for ISS configurations 1R and 9Al. The analysis is based on solving the nodal heat balance equations for ISS structure by Kutta-Merson method and analytical solutions of heat transfer equations for TCS units. The computations were performed using thermal software TERM [1,2] that will be briefly described.

Altov, Vladimir V.↗
NASA NTRS2004

Higher-Order Finite Elements for Computing Thermal Radiation

Two variants of the finite-element method have been developed for use in computational simulations of radiative transfers of heat among diffuse gray surfaces. Both variants involve the use of higher-order finite elements, across which temperatures and radiative quantities are assumed to vary according to certain approximations. In this and other applications, higher-order finite elements are used to increase (relative to classical finite elements, which are assumed to be isothermal) the accuracies of final numerical results without having to refine computational meshes excessively and thereby incur excessive computation times. One of the variants is termed the radiation sub-element (RSE) method, which, itself, is subject to a number of variations. This is the simplest and most straightforward approach to representation of spatially variable surface radiation. Any computer code that, heretofore, could model surface-to-surface radiation can incorporate the RSE method without major modifications. In the basic form of the RSE method, each finite element selected for use in computing radiative heat transfer is considered to be a parent element and is divided into sub-elements for the purpose of solving the surface-to-surface radiation-exchange problem. The sub-elements are then treated as classical finite elements; that is, they are assumed to be isothermal, and their view factors and absorbed heat fluxes are calculated accordingly. The heat fluxes absorbed by the sub-elements are then transferred back to the parent element to obtain a radiative heat flux that varies spatially across the parent element. Variants of the RSE method involve the use of polynomials to interpolate and/or extrapolate to approximate spatial variations of physical quantities. The other variant of the finite-element method is termed the integration method (IM). Unlike in the RSE methods, the parent finite elements are not subdivided into smaller elements, and neither isothermality nor other unrealistic physical conditions are assumed. Instead, the equations of radiative heat transfer are integrated numerically over the parent finite elements by use of a computationally efficient Gaussian integration scheme.

Gould, Dana C.↗
NASA NTRS2013

An Active Heater Control Concept to Meet IXO Type Mirror Module Thermal-Structural Distortion Requirement

Flight mirror assemblies (FMAs) of large telescopes, such as the International X-ray Observatory (IXO), have very stringent thermal-structural distortion requirements. The spatial temperature gradient requirement within a FMA could be as small as 0.05 C. Con ventionally, heaters and thermistors are attached to the stray light baffle (SLB), and centralized heater controllers (i.e., heater controller boards located in a large electronics box) are used. Due to the large number of heater harnesses, accommodating and routing them is extremely difficult. The total harness length/mass is very large. This innovation uses a thermally conductive pre-collimator to accommodate heaters and a distributed heater controller approach. It minimizes the harness length and mass, and reduces the problem of routing and accommodating them. Heaters and thermistors are attached to a short (4.67 cm) aluminum portion of the pre-collimator, which is thermally coupled to the SLB. Heaters, which have a very small heater power density, and thermistors are attached to the exterior of all the mirror module walls. The major portion (23.4 cm) of the pre-collimator for the middle and outer modules is made of thin, non-conductive material. It minimizes the view factors from the FMA and heated portion of the precollimator to space. It also minimizes heat conduction from one end of the FMA to the other. Small and multi-channel heater controllers, which have adjustable set points and internal redundancy, are used. They are mounted to the mechanical support structure members adjacent to each module. The IXO FMA, which is 3.3 m in diameter, is an example of a large telescope. If the heater controller boards are centralized, routing and accommodating heater harnesses is extremely difficult. This innovation has the following advantages. It minimizes the length/mass of the heater harness between the heater controllers and heater circuits. It reduces the problem of routing and accommodating the harness on the FMA. It reduces the risk of X-ray attenuation caused by the heater harness. Its adjustable set point capability eliminates the need for survival heater circuits. The operating mode heater circuits can also be used as survival heater circuits. In the non-operating mode, a lower set point is used.

Choi, Michael↗
NASA NTRS2013

Optimized Radiator Geometries for Hot Lunar Thermal Environments

The optimum radiator configuration in hot lunar thermal environments is one in which the radiator is parallel to the ground and has no view to the hot lunar surface. However, typical spacecraft configurations have limited real estate available for top-mounted radiators, resulting in a desire to use the spacecraft's vertically oriented sides. Vertically oriented, flat panel radiators will have a large view factor to the lunar surface, and thus will be subjected to significant incident lunar infrared heat. Consequently, radiator fluid temperatures will need to exceed approximately 325 K (assuming standard spacecraft radiator optical properties) in order to provide positive heat rejection at lunar noon. Such temperatures are too high for crewed spacecraft applications in which a heat pump is to be avoided. A recent study of vertically oriented radiator configurations subjected to lunar noon thermal environments led to the discovery of a novel radiator concept that yielded positive heat rejection at lower fluid temperatures. This radiator configuration, called the Intense Thermal Infrared Reflector (ITIR), has exhibited superior performance to all previously analyzed concepts in terms of heat rejection in the lunar noon thermal environment. A key benefit of ITIR is the absence of louvers or other moving parts and its simple geometry (no parabolic shapes). ITIR consists of a specularly reflective shielding surface and a diffuse radiating surface joined to form a horizontally oriented V-shape (shielding surface on top). The point of intersection of these surfaces is defined by two angles, those which define the tilt of each surface with respect to the local horizontal. The optimum set of these angles is determined on a case-by-case basis. The idea assumes minimal conductive heat transfer between shielding and radiating surfaces, and a practical design would likely stack sets of these surfaces on top of one another to reduce radiator thickness.

Ochoa, Dustin↗
NASA NTRS2016

Development, Verification and Validation of Enclosure Radiation Capabilities in the CHarring Ablator Response (CHAR) Code

With the recent development of multi-dimensional thermal protection system (TPS) material response codes including the capabilities to account for radiative heating is a requirement. This paper presents the recent efforts to implement such capabilities in the CHarring Ablator Response (CHAR) code developed at NASA's Johnson Space Center. This work also describes the different numerical methods implemented in the code to compute view factors for radiation problems involving multiple surfaces. Furthermore, verification and validation of the code's radiation capabilities are demonstrated by comparing solutions to analytical results, to other codes, and to radiant test data.

Salazar, Giovanni↗
NASA NTRS2016

Development and Verification of Enclosure Radiation Capabilities in the CHarring Ablator Response (CHAR) Code

With the recent development of multi-dimensional thermal protection system (TPS) material response codes, the capability to account for surface-to-surface radiation exchange in complex geometries is critical. This paper presents recent efforts to implement such capabilities in the CHarring Ablator Response (CHAR) code developed at NASA's Johnson Space Center. This work also describes the different numerical methods implemented in the code to compute geometric view factors for radiation problems involving multiple surfaces. Verification of the code's radiation capabilities and results of a code-to-code comparison are presented. Finally, a demonstration case of a two-dimensional ablating cavity with enclosure radiation accounting for a changing geometry is shown.

Salazar, Giovanni↗
NASA NTRS2018

Optimization of Thin-Film Solar Cells for Lunar Surface Operations

Thin-film solar cells have been in production for decades, but technology has only recently advanced enough to allow for comparable efficiencies to traditional rigid cells. Some of the benefits of thin-films, such as lighter weight and being foldable, are particularly advantageous to space applications since mass and volume are key considerations of any flight project. Using these thin-film cells in space, however, is outside of their ground-based design criteria. This requires special care to be taken in designing the power generation system of a spacecraft around a thin-film solar cell, particularly in regards to thermal management. Without the diffusion of an atmosphere to mitigate solar load, the temperature of the panels can rapidly exceed their design specification. In this paper a design solution is presented that allows for thin-film solar cells to be used in a robotic lunar lander. Due to the low thermal mass and in-plane conductivity of thin films, it is difficult to remove waste heat by any other method than radiation. On the lunar surface this means angling the arrays to increase their view factor to space, which has the negative consequence of decreasing their power generation. An optimization was developed to balance the heat rejection and power generation of the cells, using constraints on the maximum cell temperature and minimum spacecraft power requirements. The resulting solar panel angle was then used as an input to the Thermal Desktop model to verify the final panel temperatures.

Breeding, Shawn↗
NASA NTRS2019

Thermal Environment Modeling Practices for the Descent Trajectory of Lunar Landers

With the current push to send landers back to the moon, properly modeling the thermal environment for the descent is critical. Descent in this paper is described as: descending from low lunar orbit to touch down on surface. There are several challenges during this period, including: many electronics (such as battery, avionics, transponder, etc…) have higher heat loads due to higher power levels, a significant portion of the view factor to space has been blocked out by the moon making heat rejection less efficient, components that normally do not have direct line of sight to the sun may get exposure due to the lander rotating to align for descent, and thruster firing will dump more heat into the lander. All of these factors combine during the most critical phase of a lander mission make it essential that the thermal environment has been properly set up during design and analysis. This paper presents one method of setting up the thermal environment during descent in Thermal Desktop and will also include some tips and tricks.

Szerszen, Alexander↗