Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Aerodynamic Heating”

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 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Performance of LI-1542 reusable surface insulation system in a hypersonic stream

The thermal and structural performance of a large panel of LI-1542 reusable surface insulation tiles was determined by a series of cyclic heating tests using radiant lamps and aerothemal tests in the Langley 8-foot high-temperature structures tunnel. Aerothermal tests were conducted at a free-stream Mach number of 6.6, a total temperature of 1830 K, Reynolds numbers of 2.0 and 4,900,000 per meter, and dynamic pressures of 29 and 65 kPa. The results suggest that pressure gradients in gaps and flow impingement on the header walls at the end of longitudinal gaps are sources for increased gap heating. Temperatures higher than surface radiation equilibrium temperature were measured deep in gaps and at the header walls. Also, the damage tolerance of the LI-1542 tiles appears to be very high. Tile edge erosion rate was slow; could not be tolerated in a shuttle application. Tiles soaked with water and subjected to rapid depressurization and aerodynamic heating showed no visible evidence of damage.

Hunt, L. R.↗
NASA NTRS1995

Shuttle orbiter experiments: Use of an operational vehicle for advancement and validation of space systems design technologies

The NASA Orbiter Experiments (OEX) Program provided a mechanism for utilization of an operational space shuttle orbiter as a flight research vehicle, as an adjunct to its normal space transportation mission. OEX Program experiments were unique among orbiter payloads, as the research instrumentation for these experiments were carried as integral parts of the vehicle's structure, rather than being placed in the orbiter's payload bay as mission-unique cargo. On each of its first 17 flights, the Orbiter Columbia carried some type of research instrumentation. Various instrumentation systems were used to measure, in flight, the requisite parameters for determination of the orbiter aerodynamic characteristics over the entire entry flight regime and/or the aerodynamic-heating rates imposed upon the vehicle during the hypersonic portion of atmospheric entry. The data derived from this instrumentation represent benchmark hypersonic flight data heretofore unavailable for a lifting entry vehicle. The data are being used in a continual process of validation of state-of-the-art methods, both experimental and computational, for simulating/predicting the aerodynamic and aerothermal characteristics of advanced space transportation vehicles. This paper describes the OEX Program complement of research experiments, presents typical flight data obtained by these experiments, and demonstrates the utilization of these data for advancement and validation of vehicle aerothermodynamic-design tools. By example, the concept of instrumenting operational vehicles and/or spacecraft in order to perform advanced technology development and validation is demonstrated to be an effective and economical method for maturing space-systems design technologies.

Holloway, Paul F.↗
NASA NTRS1995

Documentation and archiving of the Space Shuttle wind tunnel test data base. Volume 1: Background and description

Development of the space shuttle necessitated an extensive wind tunnel test program, with the cooperation of all the major wind tunnels in the United States. The result was approximately 100,000 hours of space shuttle wind tunnel testing conducted for aerodynamics, heat transfer, and structural dynamics. The test results were converted into Chrysler DATAMAN computer program format to facilitate use by analysts, a very cost effective method of collecting the wind tunnel test results from many test facilities into one centralized location. This report provides final documentation of the space shuttle wind tunnel program. The two-volume set covers evolution of space shuttle aerodynamic configurations and gives wind tunnel test data, titles of wind tunnel data reports, sample data sets, and instructions for accessing the digital data base.

Romere, Paul O.↗
NASA NTRS1995

Documentation and archiving of the Space Shuttle wind tunnel test data base. Volume 2: User's Guide to the Archived Data Base

Development of the Space Shuttle necessitated an extensive wind tunnel test program, with the cooperation of all the major wind tunnels in the United States. The result was approximately 100,000 hours of Space Shuttle wind tunnel testing conducted for aerodynamics, heat transfer, and structural dynamics. The test results were converted into Chrysler DATAMAN computer program format to facilitate use by analysts, a very cost effective method of collecting the wind tunnel test results from many test facilities into one centralized location. This report provides final documentation of the Space Shuttle wind tunnel program. The two-volume set covers the evolution of Space Shuttle aerodynamic configurations and gives wind tunnel test data, titles of wind tunnel data reports, sample data sets, and instructions for accessing the digital data base.

Romere, Paul O.↗
NASA NTRSDate not supplied

Assessing Huygens Probe Entry, Descent, and Landing at Titan Simulation using Dragonfly Atmosphere Model

Dragonfly is a New Frontiers Program mission that will deliver a rotorcraft to Saturn's moon, Titan. This mission follows Huygens as the previous mission that successfully landed a vehicle on Titan. A flight mechanics simulation of Dragonfly's Entry, Descent, and Landing sequence has been developed using the Program to Optimize Simulated Trajectories II. The simulation incorporates several subsystem models, including aerodynamics, gravity, and mass properties, to fully capture the multi-body six degree of freedom dynamics. Among all the subsystem models that inform the Entry, Descent, and Landing dynamics, the atmosphere model of Titan is a critical component. The atmosphere model characterizes the density, temperature, pressure, and winds that the entry vehicle experiences during the descent. This impacts several aspects of the descent such as the peak heating, aerodynamics, parachute release conditions, the dynamics of the vehicle and parachutes, and the landing ellipse. In the course of developing Dragonfly, an updated model of the Titan atmosphere has been created corresponding to Dragonfly's arrival in the mid-2030s, approximately one Titan year after the Huygens mission successfully landed a probe on Titan. This work leverages previous work done to investigate Huygens EDL sequence to assess the atmosphere model developed for Dragonfly. This is done by utilizing the updated Titan atmosphere model, the Dragonfly atmosphere model, within the Huygens POST2-based flight simulation with the goal of characterizing the differences between the atmospheric models and assessing how the Dragonfly atmosphere model impacts Huygens entry dynamics.

Huygens↗
NASA NTRSDate not supplied

Assessing Huygens Probe Entry, Descent, and Landing at Titan Simulation using Dragonfly Atmosphere Model

Dragonfly is a New Frontiers Program mission that will deliver a rotorcraft to Saturn's moon, Titan. This mission follows Huygens as the previous mission that successfully landed a vehicle on Titan. A flight mechanics simulation of Dragonfly's Entry, Descent, and Landing sequence has been developed using the Program to Optimize Simulated Trajectories II. The simulation incorporates several subsystem models, including aerodynamics, gravity, and mass properties, to fully capture the multi-body six degree of freedom dynamics. Among all the subsystem models that inform the Entry, Descent, and Landing dynamics, the atmosphere model of Titan is a critical component. The atmosphere model characterizes the density, temperature, pressure, and winds that the entry vehicle experiences during the descent. This impacts several aspects of the descent such as the peak heating, aerodynamics, parachute release conditions, the dynamics of the vehicle and parachutes, and the landing ellipse. In the course of developing Dragonfly, an updated model of the Titan atmosphere has been created corresponding to Dragonfly's arrival in the mid-2030s, approximately one Titan year after the Huygens mission successfully landed a probe on Titan. This work leverages previous work done to investigate Huygens EDL sequence to assess the atmosphere model developed for Dragonfly. This is done by utilizing the updated Titan atmosphere model, the Dragonfly atmosphere model, within the Huygens POST2-based flight simulation with the goal of characterizing the differences between the atmospheric models and assessing how the Dragonfly atmosphere model impacts Huygens entry dynamics.

Entry Descent Landing↗
NASA NTRS1978

Conceptual study of hypersonic airbreathing missiles

The purpose of this paper is to report recent results of an in-house conceptual study to evaluate the performance potential and research needs of airbreathing hypersonic missiles. An alkylated-borane (noncryogenic) fueled, dual-mode, ramjet/scramjet propulsion system structured with a Rene 41 inlet and a carbon-carbon combustor was assumed along with a Lockalloy heat sink fuselage structure and beryllium wings and control surfaces. Performance for an air-launched baseline missile with a 961 pound staging weight containing a 100 pound payload indicated excellent long range cruise, moderate acceleration and high maneuverability potential. A sizing study indicates that Mach 6 cruise ranges of the order of 2500 nautical miles for payloads of 300 pounds can be achieved with moderate size missile carry weights (9000 lbs.). Aerodynamic heating analyses indicate that unprotected heat-sink structures with internal insulation are feasible for ranges of several hundred miles. For ranges of several thousands of miles a multiwall radiation shield (Inconel/titanium) was selected for protection of the internally insulated heat sink structure.

Hunt, J. L.↗
NASA NTRSDate not supplied

Phase transition energetics-based mass loss modeling of chondritic Near Earth Objects

Mitigating potentially hazardous asteroids relies on accurate knowledge of their composition and critical physical characteristics such as shape and mass. Ground-based observatories and sample return missions in recent years have focused on such characterization to track Near Earth Objects. Historically, entry speeds of asteroids or mass change due to break-up/airburst have been considered in risk assessment studies as critical parameters. To this effect, most hydrocode simulations that consider energy deposition techniques rely on gross meteor modeling with simplistic spherical shapes while neglecting heat of ablation considerations. This is the general approach for fragment-cloud, pancake and hybrid fragmentation models. Alternatively, aerothermal fragmentation modeling that accounts for meteoritic shape change have typically focused on environmental modeling to understand break-up points primarily due to radiative heating. A physics-based risk assessment could benefit by modeling the environment and material response of typical meteoroidal materials as precursors to fragmentation. A greater focus on the material response to applied environmental heating will be emphasized here. This work presents a forward analysis of thermal and material response of stony meteorites by ac-counting for mass loss and shape change effects through a boundary layer formalism. This allows for simulations starting from simplistic shapes that could eventually lead to alterations in contours due to ablation. These effects would be characterized through phase transitions of the surface, such as melting and vaporization, in response to aerodynamic heating. Coupled ablation and radiative heating effects have been found to lead to reduced heat transfer coefficients in the computation of meteor mass loss rates. This has the potential to increase ground damage footprints for hazardous impacts. The effort here quantifies the heat transfer coefficient by accounting for ablation assisted by phase transition. The melt and vaporization mediated ablation of chondritic meteors provides essential recession estimates based upon surface environments. The altered mass due to such phenomena would be presented for a H-chondrite meteoritic sample.

Pratibha Raghunandan↗
NASA NTRS1958

Heat-Transfer Measurements in Free Flight at Mach Numbers up to 14.6 on a Flat-Faced Conical Nose with a Total Angle of 29 Degrees

Skin-temperature measurements have been made at several locations on a flat-faced cone-cylinder nose which was flight tested on a fivestage rocket-propeller model to a Mach number of 14.64 and a free-stream Reynolds number of 2.0 x 10(exp 6), based on flat-face diameter, at an altitude of 66,300 feet. The copper nose had a 29 deg total-angle conical section which was 1.6 flat-face diameters long. The aerodynamic-heating rates determined from the temperature measurements reached 1,440 Btu/( sec) (sq ft) on the flat face. The heating rates near the center of the flat face agreed well at Mach numbers up to 13.6 with those obtained by a theory for laminar stagnation-point heating in equilibrium dissociated air (Avco Res. Rep. 1). At Mach numbers above 13.6, the heating rates at locations near the center of the flat face became progressively lower than stagnation-point theory and. were 29 percent lower at Mach number 14.6 at the end. of the test. The reason for this behavior of the heating on the central part of the flat face was not determined. Excluding the relatively low heating rates that occurred on the central part of the nose at the highest Mach numbers, the distribution of experimental heating along the innermost 0.79 of the flat-face radius, expressed as a percentage of stagnation-point heating, was in fair agreement with the distribution predicted by laminar theory. At a location of 0.71 radii from the stagnation point, the experimental heating was very near 130 percent of the theoretical stagnation-point rate at Mach numbers from 11 to 14.5. The experimental beating rates on the conical section of the nose were in good agreement with laminar-cone theory using the assumption of theoretical sharp-cone static pressure on the conical section.

Rumsey, Charles B↗
NASA NTRS2016

Thermal Management and Thermal Protection Systems

During my internship in the Thermal Design Branch (ES3), I contributed to two main projects: i) novel passive thermal management system for future human exploration, ii) AVCOAT undercut thermal analysis. i) As NASA prepares to further expand human and robotic presence in space, it is well known that spacecraft architectures will be challenged with unprecedented thermal environments. Future exploration activities will have the need of thermal management systems that can provide higher reliability, mass and power reduction and increased performance. In an effort to start addressing the current technical gaps the NASA Johnson Space Center Passive Thermal Discipline has engaged in technology development activities. One of these activities was done through an in-house Passive Thermal Management System (PTMS) design for a lunar lander. The proposed PTMS, functional in both microgravity and gravity environments, consists of three main components: a heat spreader, a novel hybrid wick Variable Conductance Heat Pipe (VCHP), and a radiator. The aim of this PTMS is to keep electronics on a vehicle within their temperature limits (0 and 50 C for the current design) during all mission phases including multiple lunar day/night cycles. The VCHP was tested to verify its thermal performance. I created a thermal math model using Thermal Desktop (TD) and analyzed it to predict the PTMS performance. After testing, the test data provided a means to correlate the thermal math model. This correlation took into account conduction and convection heat transfer, representing the actual benchtop test. Since this PTMS is proposed for space missions, a vacuum test will be taking place to provide confidence that the system is functional in space environments. Therefore, the model was modified to include a vacuum chamber with a liquid nitrogen shroud while taking into account conduction and radiation heat transfer. Infrared Lamps were modelled and introduced into the model to simulate the sun's rays directly impinging on the system. Heating rate of the lamps were calculated by knowing fraction of emitted energy in a wavelength interval and the filament temperature. This version of the model can be used to predict performance of the system under vacuum with extreme cold or hot conditions. Initial testing of the PTMS showed promise, and the thermal math model predicts even better performance in thermal vacuum testing. ii) Thermal Protection Systems (TPS) are required for vehicles which enter earth's atmosphere to protect from aerodynamic heating caused by the friction between the vehicle and atmospheric gases. Orion's heat shield design has two aspects which needed to be analyzed thermally: i) a small excess of adhesive used to bond the outer AVCOAT layer to the inner composite structure tends to seep from under the AVCOAT and form a small bead in between two bricks of AVCOAT, ii) a silicone rubber with different thermophysical properties than AVCOAT fills the gap between two bricks of AVCOAT. I created a thermal model using TD to determine temperature differences that are caused by these two features. To prevent false results, all TD models must be verified against something known. In this case, the TD model was correlated to CHAR, an ablation modelling software used to analyze TPS. Analyzing a node far from the concerning features, we saw that the TD model data match CHAR data, verifying the TD model. Next, the temperature of the silicone rubber as well as the bead of adhesive were analyzed to determine if they exceeded allowable temperatures. It was determined that these two features do not have a significant effect on the max temperature of the heat shield. This model can be modified to check temperatures at various locations of the heat shield where the composite thickness varies.

Hasnain, Aqib↗
NASA NTRS1985

Future fundamental combustion research for aeropropulsion systems

Physical fluid mechanics, heat transfer, and chemical kinetic processes which occur in the combustion chamber of aeropropulsion systems were investigated. With the component requirements becoming more severe for future engines, the current design methodology needs the new tools to obtain the optimum configuration in a reasonable design and development cycle. Research efforts in the last few years were encouraging but to achieve these benefits research is required into the fundamental aerothermodynamic processes of combustion. It is recommended that research continues in the areas of flame stabilization, combustor aerodynamics, heat transfer, multiphase flow and atomization, turbulent reacting flows, and chemical kinetics. Associated with each of these engineering sciences is the need for research into computational methods to accurately describe and predict these complex physical processes. Research needs in each of these areas are highlighted.

Mularz, E. J.↗
NASA NTRS1985

Future fundamental combustion research for aeropropulsion systems

Physical fluid mechanics, heat transfer, and chemical kinetic processes which occur in the combustion chamber of aeropropulsion systems were investigated. With the component requirements becoming more severe for future engines, the current design methodology needs the new tools to obtain the optimum configuration in a reasonable design and development cycle. Research efforts in the last few years were encouraging but to achieve these benefits research is required into the fundamental aerothermodynamic processes of combustion. It is recommended that research continues in the areas of flame stabilization, combustor aerodynamics, heat transfer, multiphase flow and atomization, turbulent reacting flows, and chemical kinetics. Associated with each of these engineering sciences is the need for research into computational methods to accurately describe and predict these complex physical processes. Research needs in each of these areas are highlighted.

Mularz, E. J.↗
NASA NTRS2015

Internship Abstract - Aerosciences and Flight Mechanics Intern

Mars is a hard place to land on, but my internship with NASA's Aerosciences & Flight Mechanics branch has shown me the ways in which men and women will one day land safely. I work on Mars Aerocapture, an aeroassist maneuver that reduces the fuel necessary to "capture" into Martian orbit before a descent. The spacecraft flies through the Martian atmosphere to lose energy through heating before it exits back into space, this time at a slower velocity and in orbit around Mars. Spacecraft will need to maneuver through the Martian atmosphere to accurately hit their orbit, and they will need to survive the generated heat. Engineering teams need simulation data to continue their designs, and the guidance algorithm that ensures a proper orbit insertion needs to be refined - two jobs that fell to me at the summer's start. Engineers within my branch have developed two concept aerocapture vehicles, and I run simulations on their behavior during the maneuver. I also test and refine the guidance algorithm. I spent the first few weeks familiarizing myself with the simulation software, troubleshooting various guidance bugs and writing code. Everything runs smoothly now, and I recently sent my first set of trajectory data to a Thermal Protection System group so they can incorporate it into their heat-bearing material designs. I hope to generate plenty of data in the next few weeks for various engineering groups before my internship ends mid-August. My major accomplishment so far is improving the guidance algorithm. It is a relatively new algorithm that promises higher accuracy and fuel efficiency, but it hasn't undergone extensive testing yet. I've had the opportunity to work with the principal developer - a professor at Iowa State University - to find and fix several issues. I was also assigned the task of expanding the branch's aerodynamic heating simulation software. I am excited to do this because engineers in the future will use my work to generate meaningful data and make design decisions. My internship has taught me to how to teach myself. There are no tutors, study sessions or professor office hours. When I am given an assignment I am expected to figure out how to accomplish it, and I have grown in my problem solving abilities. My summer experience has reinforced my drive to work at NASA, and I can definitely see myself working full time on the aerocapture project, or something similar.

Rangel, John↗
NASA NTRS1966

Protuberance heating test program

Results are presented of the protuberance heating test program. Four general protuberance shapes on a flat plate were tested. Presentation and evaluation of the data both on the protuberance and in the wake regions are made. The test program is an extension of the general protuberance heat transfer test. The additional series of tests was conducted to define the extent of wake heating and to assess the effects of Reynolds number variation on heating both on and around the protuberances. The protuberance models were mounted near the forward end of a six-foot instrumented test plate with stringers that simulated interstage and skirt structure of the Saturn S-4B stage. The tests were performed at Mach numbers of 2.49, 3.51, and 4.44. Reynolds numbers per foot of 3 million and 1.5 million were used for the two lower Mach numbers and 3 million for a Mach number of 4.44. The test Mach numbers simulated the Saturn S-4B flight conditions during the most severe aerodynamic heating period. The test Reynolds numbers were somewhat higher than the flight values, but lower values could not be used because of tunnel and instrumentation limitations. Oil flow runs were made on two representative models at various combinations of Mach number and Reynolds number to help define the extent of wake heating.

Sieker, W. D.↗
NASA NTRS1984

Aerothermal performance and damage tolerance of a Rene 41 metallic standoff thermal protection system at Mach 6.7

A flight-weight, metallic thermal protection system (TPS) model applicable to Earth-entry and hypersonic-cruise vehicles was subjected to multiple cycles of both radiant and aerothermal heating in order to evaluate its aerothermal performance, structural integrity, and damage tolerance. The TPS was designed for a maximum operating temperature of 2060 R and featured a shingled, corrugation-stiffened corrugated-skin heat shield of Rene 41, a nickel-base alloy. The model was subjected to 10 radiant heating tests and to 3 radiant preheat/aerothermal tests. Under radiant-heating conditions with a maximum surface temperature of 2050 R, the TPS performed as designed and limited the primary structure away from the support ribs to temperatures below 780 R. During the first attempt at aerothermal exposure, a failure in the panel-holder test fixture severely damaged the model. However, two radiant preheat/aerothermal tests were made with the damaged model to test its damage tolerance. During these tests, the damaged area did not enlarge; however, the rapidly increasing structural temperature measuring during these tests indicates that had the damaged area been exposed to aerodynamic heating for the entire trajectory, an aluminum burn-through would have occurred.

Avery, D. E.↗
NASA NTRS2000

Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Hypersonic Airbreathing Engines

The prospects for realizing a magnetohydrodynamic (MHD) bypass hypersonic airbreathing engine are examined from the standpoint of fundamental thermodynamic feasibility. The MHD-bypass engine, first proposed as part of the Russian AJAX vehicle concept, is based on the idea of redistributing energy between various stages of the propulsion system flow train. The system uses an MHD generator to extract a portion of the aerodynamic heating energy from the inlet and an MHD accelerator to reintroduce this power as kinetic energy in the exhaust stream. In this way, the combustor entrance Mach number can be limited to a specified value even as the flight Mach number increases. Thus, the fuel and air can be efficiently mixed and burned within a practical combustor length, and the flight Mach number operating envelope can be extended. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass engines using a simplified thermodynamic analysis. This cycle analysis, based on a thermally and calorically perfect gas, incorporates a coupled MHD generator-accelerator system and accounts for aerodynamic losses and thermodynamic process efficiencies in the various engin components. It is found that the flight Mach number range can be significantly extended; however, overall performance is hampered by non-isentropic losses in the MHD devices.

Litchford, R. J.↗
NASA NTRS1995

User's manual for the one-dimensional hypersonic experimental aero-thermodynamic (1DHEAT) data reduction code

A FORTRAN computer code for the reduction and analysis of experimental heat transfer data has been developed. This code can be utilized to determine heat transfer rates from surface temperature measurements made using either thin-film resistance gages or coaxial surface thermocouples. Both an analytical and a numerical finite-volume heat transfer model are implemented in this code. The analytical solution is based on a one-dimensional, semi-infinite wall thickness model with the approximation of constant substrate thermal properties, which is empirically corrected for the effects of variable thermal properties. The finite-volume solution is based on a one-dimensional, implicit discretization. The finite-volume model directly incorporates the effects of variable substrate thermal properties and does not require the semi-finite wall thickness approximation used in the analytical model. This model also includes the option of a multiple-layer substrate. Fast, accurate results can be obtained using either method. This code has been used to reduce several sets of aerodynamic heating data, of which samples are included in this report.

Hollis, Brian R.↗
NASA NTRS2002

Heat Transfer in High-Temperature Fibrous Insulation

The combined radiation/conduction heat transfer in high-porosity, high-temperature fibrous insulations was investigated experimentally and numerically. The effective thermal conductivity of fibrous insulation samples was measured over the temperature range of 300-1300 K and environmental pressure range of 1.33 x 10(exp -5)-101.32 kPa. The fibrous insulation samples tested had nominal densities of 24, 48, and 72 kilograms per cubic meter and thicknesses of 13.3, 26.6 and 39.9 millimeters. Seven samples were tested such that the applied heat flux vector was aligned with local gravity vector to eliminate natural convection as a mode of heat transfer. Two samples were tested with reverse orientation to investigate natural convection effects. It was determined that for the fibrous insulation densities and thicknesses investigated no heat transfer takes place through natural convection. A finite volume numerical model was developed to solve the governing combined radiation and conduction heat transfer equations. Various methods of modeling the gas/solid conduction interaction in fibrous insulations were investigated. The radiation heat transfer was modeled using the modified two-flux approximation assuming anisotropic scattering and gray medium. A genetic-algorithm based parameter estimation technique was utilized with this model to determine the relevant radiative properties of the fibrous insulation over the temperature range of 300-1300 K. The parameter estimation was performed by least square minimization of the difference between measured and predicted values of effective thermal conductivity at a density of 24 kilograms per cubic meters and at nominal pressures of 1.33 x 10(exp -4) and 99.98 kPa. The numerical model was validated by comparison with steady-state effective thermal conductivity measurements at other densities and pressures. The numerical model was also validated by comparison with a transient thermal test simulating reentry aerodynamic heating conditions.

Daryabeigi, Kamran↗