Revised Lunar Module Ascent Stage Thermal Model handbook
Revised Lunar Module Ascent Stage Thermal Model handbook
De Laughter, D. A.↗
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Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix B: Liquid rocket booster acoustic and thermal environments
The ascent thermal environment and propulsion acoustic sources for the Martin-Marietta Corporation designed Liquid Rocket Boosters (LRB) to be used with the Space Shuttle Orbiter and External Tank are described. Two designs were proposed: one using a pump-fed propulsion system and the other using a pressure-fed propulsion system. Both designs use LOX/RP-1 propellants, but differences in performance of the two propulsion systems produce significant differences in the proposed stage geometries, exhaust plumes, and resulting environments. The general characteristics of the two designs which are significant for environmental predictions are described. The methods of analysis and predictions for environments in acoustics, aerodynamic heating, and base heating (from exhaust plume effects) are also described. The acoustic section will compare the proposed exhaust plumes with the current SRB from the standpoint of acoustics and ignition overpressure. The sections on thermal environments will provide details of the LRB heating rates and indications of possible changes in the Orbiter and ET environments as a result of the change from SRBs to LRBs.
Source record↗
Ascent Heating Thermal Analysis on Spacecraft Adaptor Fairings
When the Crew Exploration Vehicle (CEV) is launched, the spacecraft adaptor (SA) fairings that cover the CEV service module (SM) are exposed to aero heating. Thermal analysis is performed to compute the fairing temperatures and to investigate whether the temperatures are within the material limits for nominal ascent aeroheating case. The ascent heating is analyzed by using computational fluid dynamics (CFD) and engineering codes at Marshall Space Flight Center. The aeroheating environment data used for this work is known as Thermal Environment 3 (TE3) heating data. One of the major concerns is with the SA fairings covering the CEV SM and the SM/crew launch vehicle (CLV) flange interface. The TE3 heating rate is a function of time, wall temperature, and the spatial locations. The implementation of the TE3 heating rate as boundary conditions in the thermal analysis becomes challenging. The ascent heating thermal analysis on SA fairings and SM/CLV flange interface are performed using two commercial software packages: Cullimore & Ring (C&R) Thermal Desktop (TD) 5.1 and MSC Patran 2007r1 b. TD is the pre-and post-processor for SINDA, which is a finite-difference-based solver. In TD, the geometry is built and meshed, the boundary conditions are defined, and then SINDA is used to compute temperatures. MSC Pthermal is a finite-element- based thermal solver. MSC Patran is the pre- and post-processor for Pthermal. Regarding the boundary conditions, the convection, contact resistance, and heat load can be imposed in different ways in both programs. These two software packages are used to build the thermal model for the same analysis to validate each other and show the differences in the modeling details.
Wang, Xiao Yen↗
Ascent Heating Thermal Analysis on the Spacecraft Adaptor (SA) Fairings and the Interface with the Crew Launch Vehicle (CLV)
When the crew exploration vehicle (CEV) is launched, the spacecraft adaptor (SA) fairings that cover the CEV service module (SM) are exposed to aero heating. Thermal analysis is performed to compute the fairing temperatures and to investigate whether the temperatures are within the material limits for nominal ascent aero heating case. Heating rates from Thermal Environment (TE) 3 aero heating analysis computed by engineers at Marshall Space Flight Center (MSFC) are used in the thermal analysis. Both MSC Patran 2007r1b/Pthermal and C&R Thermal Desktop 5.1/Sinda models are built to validate each other. The numerical results are also compared with those reported by Lockheed Martin (LM) and show a reasonably good agreement.
Wang, Xiao-Yen↗
Collaborative Proposal: Improving understanding of the internal structure and dynamics of deep convection using ARM observations and large eddy simulations
Recent observational and large eddy simulation (LES) modeling studies have nearly unanimously supported the view of deep cumulus convection being composed of a series of quasi-spherical bubbles of buoyant air, known as moist thermals. Despite the prevalence of moist thermals in deep convection, a comprehensive theory for the dynamics of these structures is lacking. Most current conceptual models for cumulus convection are based on canonical scaling theories for dry thermals or plumes; however, there is considerable evidence that the behavior of moist thermals differs markedly from these theories. Furthermore, the theoretical basis for most cumulus parameterizations originates from the plume conceptual model, and therefore these parameterizations are inconsistent with the real structure of moist convection. Motivated by the aforementioned knowledge gaps, this “end-to-end” research effort use theory, observations, numerical simulations, and direct improvements to the Zhang-McFarlane (ZM) convection scheme in the global climate Community Atmosphere Model (CAM) to address the following research questions: What key environmental parameters determine whether or not shallow convection will transition into deep convection, in the context of thermal-like updrafts? What factors regulate the size of thermals within cumulus updrafts? How does vertical wind shear influence thermal behavior, and as a consequence, vertical velocity and mass flux profiles and the shallow-to-deep convective transition? What are the critical processes that determine updraft vertical velocities and their connection to the vertical mass flux profile for thermal-like updrafts? Idealized LES modeling will be used in conjunction with theoretical models for the core properties of thermal-like updrafts to better understand key processes that regulate thermal ascent rates and entrainment properties. Thermal-tracking procedures will be used to characterize the behavior of thermals within the LES, and recently developed direct measures of entrainment and detrainment will be used to quantify entrainment/detrainment rates. Building from these results, we will analyze the structure of moist thermals from hemispheric range-height indicator scans taken during the Atmospheric Radiation Measurement Cloud, Aerosol, and Complex Terrain Interactions (CACTI) field campaign, and from “real case” LES of CACTI events. This combined modeling and observational analysis will provide essential validation for the existing body of research on moist thermal dynamics, which is based primarily on modeling studies. With the insight gained from the aforementioned activities, we will modify the Zhang-McFarlane convection scheme to improve its representation of updraft vertical velocity and entrainment rate profiles. These process-level changes will be tested in the Community Atmosphere Model to assess the impact on global climate simulations.
54 ENVIRONMENTAL SCIENCES↗
The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant
The following paper reports on a design study of a novel space transportation concept known as a 'NIMF' (Nuclear rocket using Indigenous Martian Fuel). The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO2. This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34,000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface.
Zubrin, Robert M.↗
The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant
A design study of a novel space transportation concept called NIMF (Nuclear rocket using Indigenous Martian Fuel) is reported. In this concept, Martian CO2 gas, which constitutes 95 percent of the atmosphere, is liquified by simple compression to about 100 psi and remains stable without refrigeration. When heated and exhausted out of a rocket nozzle, a specific impulse of about 264 s can be achieved, sufficient for flights from the surface to highly energetic orbits or from one point on the surface to any other point. The propellant acquisition system can travel with the vehicle, allowing it to refuel itself each time it lands. The concept offers unequalled potential to achieve planetwide mobility, allowing complete global access for the exploration of Mars. By eliminating the necessity of transporting ascent propellant to Mars, the NIMF can also significantly reduce the initial mass in LEO and of a manned Mars mission.
Zubrin, Robert M.↗
Effects of Planetary Thermal Structure on the Ascent and Cooling of Magma on Venus
Magellan radar images of the surface of Venus show a spatially broad distribution of volcanic features. Models of magmatic ascent processes to planetary surfaces indicate that the thermal structure of the interior significantly influences the rate of magmatic cooling and thus the amount of magma that can be transported to the surface before solidification. In order to understand which aspects of planetary thermal structure have the greatest influence on the cooling of buoyantly ascending magma, we have constructed magma cooling profiles for a plutonic ascent mechanism, and evaluated the profiles for variations in the surface and mantle temperature, surface temperature gradient, and thermal gradient curvature. Results show that, for a wide variety of thermal conditions, smaller and slower magma bodies are capable of reaching the surface on Venus compared to Earth, primarily due to the higher surface temperature of Venus. Little to no effect on the cooling and transport of magma are found to result from elevated mantle temperatures, elevation-dependent surface temperature variations, or details of the thermal gradient curvature. The enhanced tendency of magma to reach the surface on Venus may provide at least a partial explanation for the extensive spatial distribution of observed volcanism on the surface.
Sakimoto, Susan E. H.↗
Transient thermal performance of multilayer insulation systems during simulated ascent pressure decay
Transient thermal performance of multilayer insulation systems with liquid hydrogen tank during simulated Saturn 5 ascent pressure decay
Maloy, J. E.↗
Mars Sample Thermal Control During Mars Ascent and Orbit
Although NASA has no official plans at this time for a mission to return samples from Mars, the Program Formulation Office of the Mars Exploration Program sponsors ongoing mission concept studies, systems analyses, and technology investments which explore different strategies for the potential return of samples from Mars, consistent with the charter of the program and stated priorities of the science community. Maintaining the thermal integrity of collected samples would be very important. In general, samples would be collected, sealed inside tubes, and left on the surface for later retrieval. They would then be inserted into an OS (Orbiting Sample), and carried to a Mars or Solar orbit via a MAV (Mars Ascent Vehicle). Subsequently, an Earth return vehicle would rendezvous with the OS and bring it back to Earth. During ascent from Mars, the OS could serve as the nose cone of the MAV and would be subjected to significant aerodynamic heating from the Martian atmosphere. Once the OS is released from the MAV, its external surface would be exposed to potentially several years of sunlight, eclipse, planetary IR, albedo, and space. The challenge is to ensure that these samples are kept at thermally moderate conditions to preserve their integrity in these widely different environments. Various thermal techniques have been investigated to achieve sample thermal control: use of thermal protection shields and surfaces (ablative and non-ablative) to protect them from adverse exposure to ascent heating, as well combinations of thermo-optical coatings during the orbital phase. The work described herein is part of this ongoing effort & will describe the key challenges related to the thermal control of the potential Mars samples during these phases and the corresponding schemes to overcome them.
Bhandari, Pradeep↗
The effects of Venus' thermal structure on buoyant magma ascent
The recent Magellan images have revealed a broad spatial distribution of surface volcanism on Venus. Previous work in modeling the ascent of magma on both Venus and Earth has indicated that the planetary thermal structure significantly influences the magmatic cooling rates and thus the amount of magma that can be transported to the surface before solidification. In order to understand which aspects of the thermal structure have the greatest influence on the cooling of ascending magma, we have constructed magma cooling curves for both plutonic and crack buoyant ascent mechanisms, and evaluated the curves for variations in the planetary mantle temperature, thermal gradient curvature with depth, surface temperature gradient, and surface temperature. The planetary thermal structure is modeled as T/T(sub 0) = 1-tau(1-Z/Z(sub 0)(exp n), where T is the temperature, T(sub 0) is the source depth temperature, tau = 1-(T(sub s)/T(sub 0)) where T(sub s) is the planetary surface temperature, Z is the depth, Z(sub 0) is the source depth, and n is a constant that controls thermal gradient curvature with depth. The equation is used both for mathematical convenience and flexibility, as well as its fit to the thermal gradients predicted by the cooling half-space models. We assume a constant velocity buoyant ascent, body-averaged magma temperatures and properties, an initially crystal-free magma, and the same liquidus and solidus for both Venus and Earth.
Sakimoto, S. E. H.↗
Development and Qualification of Alternate Blowing Agents for Space Shuttle External Tank Thermal Protection System
The Aerospace industry has a long history of using low density polyurethane and polyurethane-modified isocyanurate foam systems as lightweight, low cost, easily processed cryogenic Thermal Protection Systems (TPS) for ascent vehicles. The Thermal Protection System of the Space Shuttle External Tank (ET) is required so that quality liquid cryogenic propellant can be supplied to the Orbiter main engines and to protect the metal structure of the tanks from becoming too hot from aerodynamic heating, hence preventing premature break-up of the tank. These foams are all blown with CFC-1 I blowing agent which has been identified by the Environmental Protection Agency (EPA) as an ozone depleting substance. CFCs will not be manufactured after 1995, Consequently, alternate blowing agent substances must be identified and implemented to assure continued ET manufacture and delivery. This paper describes the various testing performed to select and qualify HCFC-1 41 b as a near term drop-in replacement for CFC-11. Although originally intended to be a one for one substitution in the formulation, several technical issues were identified regarding material performance and processability which required both formulation changes and special processing considerations to overcome. In order to evaluate these material changes, each material was subjected to various tests to qualify them to meet the various loads imposed on them during long term storage, pre-launch operations, launch, separation and re-entry. Each material was tested for structural, thermal, aeroshear, and stress/strain loads for the various flight environments each encounters. Details of the development and qualification program and the resolution of specific problems are discussed in this paper.
Williams, Charles W.↗
Sublimator Driven Coldplate Engineering Development Unit Test Results and Development of Second Generation SDC
The Sublimator Driven Coldplate (SDC) is a unique piece of thermal control hardware that has several advantages over a traditional thermal control scheme. The principal advantage is the possible elimination of a pumped fluid loop, potentially increasing reliability and reducing complexity while saving both mass and power. Furthermore, the Integrated Sublimator Driven Coldplate (ISDC) concept couples a coolant loop with the previously described SDC hardware. This combination allows the SDC to be used as a traditional coldplate during long mission phases. The previously developed SDC technology cannot be used for long mission phases due to the fact that it requires a consumable feedwater for heat rejection. Adding a coolant loop also provides for dissimilar redundancy on the Altair Lander ascent module thermal control system, which is the target application for this technology. Tests were performed on an Engineering Development Unit at NASA s Johnson Space Center to quantify and assess the performance of the SDC. Correlated thermal math models were developed to help explain the test data. The paper also outlines the preliminary results of an ISDC concept being developed.
Stephan, Ryan A.↗
NASA Tech Briefs, May 2011
Topics covered include: 1) Method to Estimate the Dissolved Air Content in Hydraulic Fluid; 2) Method for Measuring Collimator-Pointing Sensitivity to Temperature Changes; 3) High-Temperature Thermometer Using Cr-Doped GdAlO3 Broadband Luminescence; 4)Metrology Arrangement for Measuring the Positions of Mirrors of a Submillimeter Telescope; 5) On-Wafer S-Parameter Measurements in the 325-508-GHz Band; 6) Reconfigurable Microwave Phase Delay Element for Frequency Reference and Phase-Shifter Applications; 7) High-Speed Isolation Board for Flight Hardware Testing; 8) High-Throughput, Adaptive FFT Architecture for FPGA-Based Spaceborne Data Processors; 9) 3D Orbit Visualization for Earth-Observing Missions; 10) MaROS: Web Visualization of Mars Orbiting and Landed Assets; 11) RAPID: Collaborative Commanding and Monitoring of Lunar Assets; 12) Image Segmentation, Registration, Compression, and Matching; 13) Image Calibration; 14) Rapid ISS Power Availability Simulator; 15) A Method of Strengthening Composite/Metal Joints; 16) Pre-Finishing of SiC for Optical Applications; 17) Optimization of Indium Bump Morphology for Improved Flip Chip Devices; 18) Measuring Moisture Levels in Graphite Epoxy Composite Sandwich Structures; 19) Marshall Convergent Spray Formulation Improvement for High Temperatures; 20) Real-Time Deposition Monitor for Ultrathin Conductive Films; 21) Optimized Li-Ion Electrolytes Containing Triphenyl Phosphate as a Flame-Retardant Additive; 22) Radiation-Resistant Hybrid Lotus Effect for Achieving Photoelectrocatalytic Self-Cleaning Anticontamination Coatings; 23) Improved, Low-Stress Economical Submerged Pipeline; 24) Optical Fiber Array Assemblies for Space Flight on the Lunar Reconnaissance Orbiter; 25) Local Leak Detection and Health Monitoring of Pressurized Tanks; 26) Dielectric Covered Planar Antennas at Submillimeter Wavelengths for Terahertz Imaging; 27) Automated Cryocooler Monitor and Control System; 28) Broadband Achromatic Phase Shifter for a Nulling Interferometer; 29) Super Dwarf Wheat for Growth in Confined Spaces; 30) Fine Guidance Sensing for Coronagraphic Observatories; 31) Single-Antenna Temperature- and Humidity-Sounding Microwave Receiver; 32) Multi-Wavelength, Multi-Beam, and Polarization-Sensitive Laser Transmitter for Surface Mapping; 33) Optical Communications Link to Airborne Transceiver; 34) Ascent Heating Thermal Analysis on Spacecraft Adaptor Fairings; 35) Entanglement in Self-Supervised Dynamics; 36) Prioritized LT Codes; 37) Fast Image Texture Classification Using Decision Trees; 38) Constraint Embedding Technique for Multibody System Dynamics; 39) Improved Systematic Pointing Error Model for the DSN Antennas; 40) Observability and Estimation of Distributed Space Systems via Local Information-Exchange Networks; 41) More-Accurate Model of Flows in Rocket Injectors; 42) In-Orbit Instrument-Pointing Calibration Using the Moon as a Target; 43) Reliability of Ceramic Column Grid Array Interconnect Packages Under Extreme Temperatures; 44) Six Degrees-of-Freedom Ascent Control for Small-Body Touch and Go; and 45) Optical-Path-Difference Linear Mechanism for the Panchromatic Fourier Transform Spectrometer.
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Polymeric Materials Application and Testing for Space Missions
The EM41 Non-metallics polymers team performs polymeric materials screening, characterization, development, qualification, verification and validation testing for flight hardware and in-space applications. They perform insight, testing and application of polymeric materials for launch vehicles and space environments. Personnel have a deep understanding of the technologies associated with cryogenic insulation, high temperature thermal protection materials, paints, primers, coatings, and adhesive materials. Numerous TPS material characterization tests can also be performed by the Lab. The team has both internal and partner test capabilities to subject polymeric materials to Aerospace and Space environments Material properties, response and performance are tested within environments from the launch pad, through the launch environment and into space. A primary role of the team is to select materials and organize test regimen then analyze test results. Test regiment are established to confirm the materials meet the performance requirements levied by customers within the environments that are expected for the mission.
Thermal↗
Optical Properties of Thermal Control Coatings After Weathering, Simulated Ascent Heating, and Simulated Space Radiation Exposure
Spacecraft radiators reject heat to their surroundings and coatings play an important role in this heat rejection. The coatings provide the combined optical properties of low solar absorptance and high infrared emittance. The coatings are applied to the radiator panel in a number of ways, including conventional spraying, plasma spraying, or as an applique. Not designed for a terrestrial weathering environment, the durability of spacecraft paints, coatings, and appliques upon exposure to weathering and subsequent exposure to ascent heating, solar wind, and ultraviolet radiation was studied. In addition to traditional aluminum panels, new isocyanate ester composite panels were exposed for a total of 90 days at the Atmospheric Exposure Site of Kennedy Space Center's (KSC) Beach Corrosion Facility for the purpose of identifying their durability to weathering. Selected panel coupons were subsequently exposed to simulated ascent heating, solar wind, and vacuum ultraviolet (UV) radiation to identify the effect of a simulated space environment on as-weathered surfaces. Optical properties and adhesion testing were used to document the durability of the paints, coatings, and appliques.
Jaworske, Donald A.↗
Lithospheric thinning associated with rifting in East Africa
The rate of ascent of the thermal perturbation of eruptive rock types in the Kenya Dome region can be calculated for depths of origin ranging from greater than 170 km to the lower crust (25-30 km) as combined with dated occurrences. An equation is derived and solved iteratively for ascent velocity at 10 km depth intervals from 150-30 km, assuming a thermal diffusivity of 1 sq mm/sec. For radii of 50 and 100 km and constant gradient increases within the range 55-60 C/km over the lithosphere-asthenosphere boundary (LAB) ascent depth range, the calculated rate of upwelling corresponds well with the results of a simple thermal model of lithospheric thinning. A 1 C/km change in gradient results approximately in a 0.55 Myr change in ascent time from 150 to 30 km, due to the increasing differential between the geotherm and the solidus with decreasing depths and the increasing rate of heat loss to the surface. It is suggested by the temporal-spatial-compositional relations of eruptives in the Kenya Dome region that there is a shallowing of magma source regions with time, implying encroachment of a thermal anomaly.
Wendlandt, R. F.↗