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Pulling device consisting of structural tube, connecting rod, spring-loaded nuts, loading rod, heating element, and three bulkheads fractures tensile specimens. Alternate heating and cooling increases tensile loading by increments until fracturing occurs. Load cell or strain gage, applied to pulling rod, determines forces applied.
The subject orifice pulse tube cooler employs dual opposed compressor pistons, clearance seals, moving coil linear motors, and flexible bearings cooling capacity is 100 watts at 200 K. A matrix of optimizations explored load temperatures and cooling capacities of 50 K, 100 K, 200 K, 1 watt, 10 watts, and 100 watts. The optimization process resulted in a design with zero second entrance flow and a non-resonant pulse tube. Topics include performance results, mechanical design, and development problems and solutions. The development seeks to use orifice pulse tube mechanical simplicity to achieve a computer cooler with acceptable efficiency, low materials cost, low labor cost, long life, high reliability, low vibration, operation at the standard 60 Hz U.S. grid frequency, and control through simple drive and control electronics.
The scientific balloon program at NASA offers an exciting and open area of opportunity for testing new technologies and for conducting meaningful experimentation at a fraction of the cost of a space mission. This paper outlines a simple thermal model developed and employed for the Primordial Inflation Polarization ExplorER (PIPER). The sub-orbital environment that PIPER operates in hosts an interesting mix of atmospheric and space thermal challenges. The work done was to mitigate thermal loads and passively cool the payload's exterior mounted electronics at an altitude of 36.6 km. This was done by characterizing the thermal environment and then designing solutions for the heat loads through a combined radiation and conduction passive cooling radiator system thermal model. Despite the simplicity and subsequent limitations of the model, as well as some unexpected payload operational events, the model produced results between 0.31% and 11.8% difference between the predicted values and measured average temperatures. From these results, the model was able to successfully provide estimates for the electronics temperatures. Additional flights will be needed to eliminate unknowns encountered in this flight in order to further refine the model for higher accuracy.
A procedure was developed to evaluate the cost effectiveness of combining an absorption cycle chiller with a solar energy system. A basic assumption of the procedure is that a solar energy system exists for meeting the heating load of the building, and that the building must be cooled. The decision to be made is to either cool the building with a conventional vapor compression cycle chiller or to use the existing solar energy system to provide a heat input to the absorption chiller. Two methods of meeting the cooling load not supplied by solar energy were considered. In the first method, heat is supplied to the absorption chiller by a boiler using fossil fuel. In the second method, the load not met by solar energy is net by a conventional vapor compression chiller. In addition, the procedure can consider waste heat as another form of auxiliary energy. Commercial applications of solar cooling with an absorption chiller were found to be more cost effective than the residential applications. In general, it was found that the larger the chiller, the more economically feasible it would be. Also, it was found that a conventional vapor compression chiller is a viable alternative for the auxiliary cooling source, especially for the larger chillers. The results of the analysis gives a relative rating of the sites considered as to their economic feasibility of solar cooling.
The electronic link connecting cryogenically cooled radiation detectors to data acquisition and signal processing electronics at higher temperatures contributes significantly to the total heat load on spacecraft cooling systems that use combined mechanical and cryogenic liquid cooling. Using high transition temperature superconductors for this link has been proposed to increase the lifetime of space missions. Herein, several YBCO (YBa2Cu3O7) superconductor-substrate combinations were examined and total heat loads were compared to manganin wire technology in current use. Using numerical solutions to the heat-flow equations, it is shown that replacing manganin technology with YBCO thick film technology can extend a 7-year mission by up to 1 year.
The Orion spacecraft has recently completed Artemis II, the first crewed mission around the moon in 54 years. This test flight demonstrated vehicle capability as the foundation for all future missions under the Artemis program. As a major system of Orion, the Active Thermal Control System (ATCS) successfully managed crew and vehicle heat loads to provide adequate cooling and thermal comfort for the duration of the Artemis II mission. Thermal control on the Orion vehicle is managed through two redundant cooling loops in the Crew Module (CM) that absorb heat loads from the cabin and Air Revitalization System (ARS) heat exchangers, the Liquid Cooling Garment (LCG) heat exchanger, and the avionics cold plates, and then transfer those heat loads to the Service Module (SM) radiator loops and/or the Phase Change Material (PCM) and ammonia boilers when required through thermal configuration or mission phase. During Artemis II, Orion did not necessitate any supplemental cooling from ammonia boilers until service module separation, allowing the maximum availability of ammonia cooling prior to entry. This presentation will provide an overview of the Orion ATCS major subsystems, highlight the performance during key mission phases, and comment on the applicability of the observed thermal response to future Artemis missions.
The solar energy research is reported including climatic data, architectural data, heating and cooling equipment, thermal loads, and economic data. Lists of data sources presented include: selected data sources for solar energy heating and cooling; bibliography of solar energy, and other energy sources; sources for manufacturing and sales, solar energy collectors; and solar energy heating and cooling projects.
In current and advanced gas turbine engines, increased speeds, pressures and temperatures are used to reduce specific fuel consumption and increase thrust/weight ratios. Hence, the turbine airfoils are subjected to increased heat loads escalating the cooling requirements to satisfy life goals. The efficient use of cooling air requires that the details of local geometry and flow conditions be adequately modeled to predict local heat loads and the corresponding heat transfer coefficients. The objective of this program is to develop a heat transfer and pressure drop data base, computational fluid dynamic techniques and correlations for multi-pass rotating coolant passages with and without flow turbulators. The experimental effort is focused on the simulation of configurations and conditions expected in the blades of advanced aircraft high pressure turbines. With the use of this data base, the effects of Coriolis and buoyancy forces on the coolant side flow can be included in the design of turbine blades.
The objective of this study is to investigate the effect of nozzle out-of-roundness on the transient startup side loads. The out-of-roundness could be the result of asymmetric loads induced by hardware attached to the nozzle, asymmetric internal stresses induced by previous tests and/or deformation, such as creep, from previous tests. The rocket engine studied encompasses a regeneratively cooled thrust chamber and a film cooled nozzle extension with film coolant distributed from a turbine exhaust manifold. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet history based on an engine system simulation. Transient startup computations were performed with the out-of-roundness achieved by four degrees of ovalization of the nozzle: one perfectly round, one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The computed side load physics caused by the nozzle out-of-roundness and its effect on nozzle side load are reported and discussed.
The objective of this study is to investigate the effect of nozzle out-of-roundness on the transient startup side loads at a high altitude, with an anchored computational methodology. The out-of-roundness could be the result of asymmetric loads induced by hardware attached to the nozzle, asymmetric internal stresses induced by previous tests, and deformation, such as creep, from previous tests. The rocket engine studied encompasses a regeneratively cooled thrust chamber and a film cooled nozzle extension with film coolant distributed from a turbine exhaust manifold. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet history based on an engine system simulation. Transient startup computations were performed with the out-of-roundness achieved by four different degrees of ovalization: one perfectly round, one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation-line-jump is the peak side load physics for the round, slightly our-of-round, and more out-of-round cases, and the peak side load increases as the degree of out-of-roundness increases. For the significantly out-of-round nozzle, however, the peak side load reduces to comparable to that of the round nozzle and the separation line jump is not the peak side load physics. The counter-intuitive result of the significantly out-of-round case is found to be related to a side force reduction mechanism that splits the effect of the separation-line-jump into two parts, not only in the circumferential direction and most importantly in time.
There was interest in understanding the impact of out-of-round nozzle extension on the nozzle side load during transient startup operations. The out-of-round nozzle extension could be the result of asymmetric internal stresses, deformation induced by previous tests, and asymmetric loads induced by hardware attached to the nozzle. The objective of this study was therefore to computationally investigate the effect of out-of-round nozzle extension on the nozzle side loads during an engine startup transient. The rocket engine studied encompasses a regeneratively cooled chamber and nozzle, along with a film cooled nozzle extension. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and transient inlet boundary flow properties derived from an engine system simulation. Six three-dimensional cases were performed with the out-of-roundness achieved by three different degrees of ovalization, elongated on lateral y and z axes: one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation line jump was the primary source of the peak side loads. Comparing to the peak side load of the perfectly round nozzle, the peak side loads increased for the slightly and more ovalized nozzle extensions, and either increased or decreased for the two significantly ovalized nozzle extensions. A theory based on the counteraction of the flow destabilizing effect of an exacerbated asymmetrical flow caused by a lower degree of ovalization, and the flow stabilizing effect of a more symmetrical flow, created also by ovalization, is presented to explain the observations obtained in this effort.
Three-dimensional numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, and pressure-based computational fluid dynamics formulation, and a simulated inlet condition based on a system calculation. Finite-rate chemistry was used throughout the study so that combustion effect is always included, and the effect of wall cooling on side load physics is studied. The side load physics captured include the afterburning wave, transition from free- shock to restricted-shock separation, and lip Lambda shock oscillation. With the adiabatic nozzle, free-shock separation reappears after the transition from free-shock separation to restricted-shock separation, and the subsequent flow pattern of the simultaneous free-shock and restricted-shock separations creates a very asymmetric Mach disk flow. With the cooled nozzle, the more symmetric restricted-shock separation persisted throughout the start-up transient after the transition, leading to an overall lower side load than that of the adiabatic nozzle. The tepee structures corresponding to the maximum side load were addressed.
The characteristics of a cryogenic cooling system for the Pollution Monitoring Satellite (PMS) are discussed. Studies were conducted to make the following determinations: (1) the characteristics and use of proven and state-of-the-art cryogenic cooling systems for six specified ranges of performance, (2) the system most applicable for each of the six cooling categories, and (3) conceptual designs for candidate system for each of the six representative cooling categories. The six cooling categories of electrical loads are defined. The desired mission life for the cooling system is two years with both continuous and intermittent operating conditions.
Parametric studies were conducted to investigate the relative merits of construction materials, coolants, and cooled panel concepts for internal convective cooling systems applied to airframe structures of hydrogen-fueled hypersonic aircraft. These parametric studies were then used as a means of comparing various cooled structural arrangements for a hypersonic transport and a hypersonic research airplane. The cooled airplane studies emphasized weight aspects as related to the choice of materials, structural arrangements, structural temperatures, and matching of the cooling system heat load to the available hydrogen fuel-flow heat sink. Consideration was given to reliability and to fatigue and fracture aspects, as well. Even when auxiliary thermal protection system items such as heat shielding, insulation, and excess hydrogen for cooling are considered the more attractive actively cooled airframe concepts indicated potential payload increases of from 40 percent to over 100 percent for the hypersonic transport as compared to the results of previous studies of the same vehicle configuration with an uncooled airframe.
Future large infrared space telescopes will require cooling to 4K to achieve background limited performance for submillimeter wavelengths. These observatories will require lifetimes of many years and will have relatively large cooling requirements making stored helium dewars impractical. We have designed and are building an adiabatic demagnetization refrigerator (ADR) for use in cooling relatively large loads (10- 100 mW) at 4K and rejecting that heat to a cryocooler operating at 1 OK. Cryocoolers below 1 OK have poor thermodynamic efficiency and ADRs can operate in this temperature range with an efficiency of 75% of Carnot or better. Overall, this can save as much as 2/3 of the input power required to operate a 4K cryocooler. The ADR magnet consists of 8 short coils wired in series and arranged in a toroid to provide self shielding of its magnetic field. This will save mass (about 30% of the mass or about 1.5 kg in our small version, higher percentages in higher cooling power, larger versions) that would have been used for passive or active shields in an ordinary solenoid. The toroid has a 100 mm outer diameter and will produce an approximately 3T average field. In the initial demonstration model the toroid coils will be wound with ordinary NbTi wire and operated at 4K. A second version will then use Nb3Sn wire to provide complete 10K operation. As a refrigerant for this temperature range we will use either GdLiF4 or GdF3 crystals, pending tests of these crystals' cooling capacity per field and thermal conductance. Preliminary indications are that these materials are superior to GGG. We will use gas gap heat switches to alternately connect the toroid to the cold load and the warm heat sink. A small continuous stage will maintain the cold end at 4K while the main toroid is recycled.
Gear tooth average and instantaneous surface temperatures were measured with a fast response infrared radiometric microscope, while operating at arious speeds, loads and oil jet pressures. Increased oil jet pressure had a significant effect on both average and peak surface temperatures at all test conditions, increasing the speed at constant load and increasing the load at constant speed causes a significant rise in average and peak surface temperatures of gear teeth. A gear tooth temperature analysis was conducted by a finite element method combined with a calculated heat input and oil jet impingment depth with estimated heat transfer coefficients based on the experimental data. It is concluded that oil jet pressures required for adequate cooling at high load and speed conditions must be high enough to get full penetration depth of the teeth.
Elastic plastic stress strain states in cooled turbine blade airfoils were calculated by three methods for the initial takeoff transient of an advanced technology aircraft engine. The three analytical methods compared were a three dimensional elastic plastic, finite element analysis, a three dimensional, elastic, finite element analysis, and a one dimensional, elastic plastic, beam theory analysis. Structural analyses were performed for eight cases involving different combinations of mechanical and thermal loading on impingement cooled airfoils with and without leading edge film cooling holes. The von Mises effective total strains at maximum takeoff computed from the elastic and elastic plastic finite element analyses agreed with 9 percent for rotating airfoils and 28 percent for stationary airfoils with the elastic results on the conservative side.
An analytical evaluation of a new typ An analytical evaluation of a new type of air-cooled turbine-rotor-blade design, based on the principle of submerging the load-carrying element in cooling air within a thin high-temperature sheel, indicates that this principle of blade design permits the load carrying element to be operated at considerably lower temperature than that of the enveloping shell. Comparison with an air-cooled shell-supported air-cooled blade has greater potentiality to withstand increased stresses that can be anticipated in future engines.