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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Integral Channel Nozzles and Heat Exchangers using Additive Manufacturing Directed Energy Deposition NASA HR-1 Alloy

Heat exchangers for use in propulsion applications are very critical components because they must be efficient, compact and light and often operate with working fluids at extreme temperatures or pressures or both. Various components and systems use heat exchangers such as combustion chambers of gas turbines and internal combustion engines, fuel cells (air supply and thermal management), electric batteries (thermal management), evaporators and recuperators of waste-heat-to-power systems, and rocket engines. Even if the results are more generally applicable, the heat exchangers applications to which this study is more closely related are regeneratively cooled rocket nozzles and chambers, and repressurization systems for the launch vehicles. These components are often thin-walled and contain pressurized fluids, like propellants at cryogenic or elevated temperatures. Given that the environments that these propulsion components must endure are challenging, the manufacturing to meet these specifications often require long lead times due to specialty processes and unique tooling associated with the combined thin-wall integral channel and large-scale structures. Additive manufacturing (AM) offers programmatic advantages for reduction in processing time and cost in addition to various technical advantages, including the possibility to achieve enhanced hardware complexity targeted to superior performance, part consolidation, and the capability of processing of novel alloys. While AM is already being utilized for heat exchanger components in propulsion applications, almost all these AM components are made by means of Laser Powder Bed Fusion (L-PBF). L-PBF allows for fine features but is rather limited with respect to the overall size of the components that can be manufactured. Recent developments are maturing the Laser Powder Directed Energy Deposition (LP-DED) process which may be used, for example, to make integral channel thin-wall regeneratively-cooled rocket nozzles with diameters greater than 1 m. This paper highlights some integral channel heat exchanger demonstrator hardware applications of LP-DED, as well as the characterization of this process in combination with the use of the NASA HR-1 alloy. To properly utilize LP-DED for heat exchanger manufacturing, various aspects are being characterized such as geometry limitations, measurement of surface texture and geometric angled surfaces, surface enhancements for internal channels, and material evaluation. NASA HR-1 (FeNi-Cr) is a high strength hydrogen resistant superalloy developed for use in aerospace applications, such as heat exchangers. Some aspects and considerations about the design of heat exchangers are summarized together with data relevant to LP-DED manufacturing in combination with the NASA HR-1 alloy. Microchannels were successful deposited down to 2.54 mm and 1 mm wall thickness, wall angles of 30°, both with high reproducibility. It was also found that the areal surface roughness is highly dependent on the size of the powder feedstock used for deposition. The characterization of these LP-DED features is critical for fluid flow and heat transfer predictions as it can be exploited to enhance heat transfer at the cost of increased pressure drop.

additive manufacturing↗
NASA NTRS1989

Space Power Free-Piston Stirling Engine Scaling Study

The design feasibility study is documented of a single cylinder, free piston Stirling engine/linear alternator (FPSE/LA) power module generating 150 kW-electric (kW sub e), and the determination of the module's maximum feasible power level. The power module configuration was specified to be a single cylinder (single piston, single displacer) FPSE/LA, with tuning capacitors if required. The design requirements were as follows: (1) Maximum electrical power output; (2) Power module thermal efficiency equal to or greater than 20 percent at a specific mass of 5 to 8 kg/kW(sub e); (3) Heater wall temperature/cooler wall temperature = 1050 K/525 K; (4) Sodium heat-pipe heat transport system, pumped loop NaK (sodium-potassium eutectic mixture) rejection system; (5) Maximum power module vibration amplitude = 0.0038 cm; and (6) Design life = 7 years (60,000 hr). The results show that a single cylinder FPSE/LA is capable of meeting program goals and has attractive scaling attributes over the power range from 25 to 150 kW(sub e). Scaling beyond the 150 kW(sub e) power level, the power module efficiency falls and the power module specific mass reaches 10 kg/kW(sub e) at a power output of 500 kW(sub e). A discussion of scaling rules for the engine, alternator, and heat transport systems is presented, along with a detailed description of the conceptual design of a 150 kW(sub e) power module that meets the requirements. Included is a discussion of the design of a dynamic balance system. A parametric study of power module performance conducted over the power output range of 25 to 150 kW(sub e) for temperature ratios of 1.7, 2.0, 2.5, and 3.0 is presented and discussed. The results show that as the temperature ratio decreases, the efficiency falls and specific mass increases. At a temperature ratio of 1.7, the 150 kW(sub e) power module cannot satisfy both efficiency and specific mass goals. As the power level increases from 25 to 150 kW(sub e) at a fixed temperature ratio, power module efficiency is seen to increase slightly, but at the expense of increased specific mass. An empirical equation relating power module thermal efficiency as a function of power module specific mass, power output, and temperature ratio is developed. Alternative configurations to the single cylinder, direct coupled linear alternator approach are also evaluated, but are shown to have technical drawbacks that lessen their attractiveness. The dynamic balance assembly mass (moving mass and structure) represents 20 to 30 percent of the total single cylinder power module mass. Joining two modules in a balanced opposed configuration eliminates the need for the balancer, and a hot end junction can be made without significant addition of structural mass. Recommendations are made for evaluation of advanced heat pipe concepts, tests of radial flow heat exchangers, and evaluation of high temperature alternator materials.

Jones, D.↗
NASA NTRS2002

Pulse thermal loop

A pulse thermal loop heat transfer system includes a means to use pressure rises in a pair of evaporators to circulate a heat transfer fluid. The system includes one or more valves that iteratively, alternately couple the outlets the evaporators to the condenser. While flow proceeds from one of the evaporators to the condenser, heating creates a pressure rise in the other evaporator, which has its outlet blocked to prevent fluid from exiting the other evaporator. When the flow path is reconfigured to allow flow from the other evaporator to the condenser, the pressure in the other evaporator is used to circulate a pulse of fluid through the system. The reconfiguring of the flow path, by actuating or otherwise changing the configuration of the one or more valves, may be triggered when a predetermined pressure difference between the evaporators is reached.

Weislogel, Mark M.↗
NASA NTRS1992

Thermal control systems for low-temperature heat rejection on a lunar base

In this report, Rankine-cycle heat pumps and absorption heat pumps (ammonia-water and lithium bromide-water) have been analyzed and optimized for a lunar base cooling load of 100 kW. For the Rankine cycle, a search of several commonly used commercial refrigerants provided R11 and R717 as possible working fluids. Hence, the Rankine-cycle analysis has been performed for both R11 and R717. Two different configurations were considered for the system--one in which the heat pump is directly connected to the rejection loop and another in which a heat exchanger connects the heat pump to the rejection loop. For a marginal increase in mass, the decoupling of the rejection loop and the radiator from the heat pump provides greater reliability of the system and better control. Hence, the decoupled system is the configuration of choice. The optimal TCS mass for a 100 kW cooling load at 270 K was 5940 kg at a radiator temperature of 362 K. R11 was the working fluid in the heat pump, and R717 was the transport fluid in the rejection loop. Two TCS's based on an absorption-cycle heat pump were considered, one with an ammonia-water mixture and the other with a lithium bromide-water mixture as the working fluid. A complete cycle analysis was performed for these systems. The system components were approximated as heat exchangers with no internal pressure drop for the mass estimate. This simple approach underpredicts the mass of the systems, but is a good 'optimistic' first approximation to the TCS mass in the absence of reliable component mass data. The mass estimates of the two systems reveal that, in spite of this optimistic estimate, the absorption heat pumps are not competitive with the Rankine-cycle heat pumps. Future work at the systems level will involve similar analyses for the Brayton- and Stirling-cycle heat pumps. The analyses will also consider the operation of the pump under partial-load conditions. On the component level, a capillary evaporator will be designed, built, and tested in order to investigate its suitability in lunar base TCS and microgravity two-phase applications.

Sridhar, K. R.↗
NASA NTRS2019

Thermal Model of a Zero Boil Off System for the Nuclear Thermal Propulsion System

NASA is currently developing an updated concept for a nuclear thermal propulsion (NTP) system. To enable this concept, efficient thermal insulation and cryocooler heat exchanger systems are required to eliminate boil-off of propellant. This paper presents the results of a thermal model used to assess the feasibility of using active cooling with a tube-on-tank heat exchanger configuration for the inline tank of the NTP system. Results show that: cryocooler working fluid temperature and mass flow rate can be adjusted to achieve zero boil off (ZBO) with broad area cooling, over-sizing the cryocooler lift directly translates into a reduction in tank pressure, and broad area cooling may still maintain ZBO despite the reduced heat transfer between tank wall and propellant that is expected in reduced gravity.

zero boil off↗
NASA NTRS1992

Analytical and experimental studies of heat pipe radiation cooling of hypersonic propulsion systems

Analytical and experimental studies were completed to assess the feasibility of using high-temperature heat pipes to cool hypersonic engine components. This new approach involves using heat pipes to transport heat away from the combustor, nozzle, or inlet regions, and to reject it to the environment by thermal radiation from an external heat pipe nacelle. For propulsion systems using heat pipe radiation cooling (HPRC), it is possible to continue to use hydrocarbon fuels into the Mach 4 to Mach 6 speed range, thereby enhancing the economic attractiveness of commercial or military hypersonic flight. In the second-phase feasibility program recently completed, it is found that heat loads produced by considering both convection and radiation heat transfer from the combustion gas can be handled with HPRC design modifications. The application of thermal insulation to ramburner and nozzle walls was also found to reduce the heat load by about one-half and to reduce peak HPRC system temperatures to below 2700 F. In addition, the operation of HPRC at cruise conditions of around Mach 4.5 and at an altitude of 90,000 ft lowers the peak hot-section temperatures to around 2800 F. An HPRC heat pipe was successfully fabricated and tested at Mach 5 conditions of heat flux, heat load, and temperature.

Martin, R. A.↗
NASA NTRS2000

Practical Considerations of Waste Heat Reuse for a Mars Mission Advanced Life Support System

Energy conservation is a key issue in design optimization of Advanced Life Support Systems (ALSS) for long-term space missions. By considering designs for conservation at the system level, energy saving opportunities arise that would otherwise go unnoticed. This paper builds on a steady-state investigation of system-level waste heat reuse in an ALSS with a low degree of crop growth for a Mars mission. In past studies, such a system has been defined in terms of technology types, hot and cold stream identification and stream energy content. The maximum steady-state potential for power and cooling savings within the system was computed via the Pinch Method. In this paper, several practical issues are considered for achieving a pragmatic estimate of total system savings in terms of equivalent system mass (ESM), rather than savings solely in terms of power and cooling. In this paper, more realistic ESM savings are computed by considering heat transfer inefficiencies during material transfer. An estimate of the steady-state mass, volume and crewtime requirements associated with heat exchange equipment is made by considering heat exchange equipment material type and configuration, stream flow characteristics and associated energy losses during the heat exchange process. Also, previously estimated power and cooling savings are adjusted to reflect the impact of such energy losses. This paper goes one step further than the traditional Pinch Method of considering waste heat reuse in heat exchangers to include ESM savings that occur with direct reuse of a stream. For example, rather than exchanging heat between crop growth lamp cooling air and air going to a clothes dryer, air used to cool crop lamps might be reused directly for clothes drying purposes. When thermodynamically feasible, such an approach may increase ESM savings by minimizing the mass, volume and crewtime requirements associated with stream routing equipment.

Levri, Julie↗
NASA NTRS2002

Passive and Active Microwave Remote Sensing of Precipitation and Latent Heating Distributions in the Tropics from TRMM

Passive and active microwave remote sensing data are analyzed to identify signatures of precipitation and vertical motion in tropical convection. A database of cloud/radiative model simulations is used to quantify surface rain rates and latent heating profiles that are consistent with these signatures. At satellite footprint-scale (approximately 10 km), rain rate and latent heating estimates are subject to significant random errors, but by averaging the estimates in space and time, random errors are substantially reduced, Bias errors have been minimized by improving the microphysics in the supporting cloud/radiative model simulations, and by imposing a consistent definition of remotely-sensed and model-simulated convective/stratiform rain coverage. Remotely-sensed precipitation and latent heating distributions in the tropics are derived from Tropical Rainfall Measuring Mission (TRMM) and Special Sensor Microwave/ Imager (SSM/ I) sensor data. The prototype Version 6 TRMM passive microwave algorithm typically yields average heating profiles with maxima between 6 and 7 km altitude for organized mesoscale convective systems. Retrieved heating profiles for individual convective systems are compared to coincident estimates based upon a combination of dual-Doppler radar and rawinsonde data. Also, large-scale latent heating distributions are compared to estimates derived from a simpler technique that utilizes observations of surface rain rate and stratiform rain proportion to infer vertical heating structure. Results of these tests will be presented at the conference.

Olson, William S.↗
NASA NTRS1985

Study of advanced fuel system concepts for commercial aircraft

An analytical study was performed in order to assess relative performance and economic factors involved with alternative advanced fuel systems for future commercial aircraft operating with broadened property fuels. The DC-10-30 wide-body tri-jet aircraft and the CF6-8OX engine were used as a baseline design for the study. Three advanced systems were considered and were specifically aimed at addressing freezing point, thermal stability and lubricity fuel properties. Actual DC-10-30 routes and flight profiles were simulated by computer modeling and resulted in prediction of aircraft and engine fuel system temperatures during a nominal flight and during statistical one-day-per-year cold and hot flights. Emergency conditions were also evaluated. Fuel consumption and weight and power extraction results were obtained. An economic analysis was performed for new aircraft and systems. Advanced system means for fuel tank heating included fuel recirculation loops using engine lube heat and generator heat. Environmental control system bleed air heat was used for tank heating in a water recirculation loop. The results showed that fundamentally all of the three advanced systems are feasible but vary in their degree of compatibility with broadened-property fuel.

Coffinberry, G. A.↗
NASA NTRS1993

Thermal control systems for low-temperature heat rejection on a lunar base

One of the important issues in the design of a lunar base is the thermal control system (TCS) used to reject low-temperature heat from the base. The TCS ensures that the base and the components inside are maintained within an acceptable temperature range. The temperature of the lunar surface peaks at 400 K during the 336-hour lunar day. Under these circumstances, direct dissipation of waste heat from the lunar base using passive radiators would be impractical. Thermal control systems based on thermal storage, shaded radiators, and heat pumps have been proposed. Based on proven technology, innovation, realistic complexity, reliability, and near-term applicability, a heat pump-based TCS was selected as a candidate for early missions. In this report, Rankine-cycle heat pumps and absorption heat pumps (ammonia water and lithium bromide-water) have been analyzed and optimized for a lunar base cooling load of 100 kW.

Sridhar, K. R.↗
NASA NTRS2001

Thermal Technology Development Activities at the Goddard Space Flight Center: 2001

Activities include the development of two phase systems which are composed of 1) heat pipes and variable conductance heat pipes, 2)capillary pumped loops, 3) loop heat pipes, 4) vapor compression systems (heat pumps), 5) phase change materials. Also in the development phase are variable emittance surfaces, advanced coatings, high conductivity materials, and electrohydrodynamic (EHD) thermal control systems.

Butler, Dan↗
NASA NTRS1972

Survey of aircraft electrical power systems

Areas investigated include: (1) load analysis; (2) power distribution, conversion techniques and generation; (3) design criteria and performance capabilities of hydraulic and pneumatic systems; (4) system control and protection methods; (5) component and heat transfer systems cooling; and (6) electrical system reliability.

Lee, C. H.↗