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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.
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Tradeoffs between the different kernel methods and performance relative to more traditional methods are discussed.
The film depicts two tests of a flat roof, conical inlet canopy parachute. The first test is a series of wind tunnel trials with a flat circular ribbon roof of 22 percent porosity. The second test is a single series of wind tunnel trials with a flat circular ribbon roof of 25 percent porosity. Variables for both trials include Mach number, dynamic pressure, longitudinal separation distances (x/d), and drag coefficient C(sub d).
Multiple wind tunnel test trials were conducted on a 30 degree conical ribbon parachute with porosities of 30, 27, and 24 percent. Variables were Mach number, dynamic pressure, towline length, and coefficient of drag. A Rotofoil parachute having a porosity of approximately 24 percent was tested, but failed after about 30 seconds of operation at a Mach number of 1.8 All of the parachutes had a nominal diameter and shroud line length of 10 inches. Drag coefficients were based on the area of a circle having a diameter two-thirds of the nominal parachute diameter.
Boundary layer transition over axisymmetric bodies at non-zero angle of attack in supersonic flow is numerically investigated as part of joint research between the National Aeronautics and Space Administration (NASA) and Japan Aerospace Exploration Agency (JAXA). Transition over four axisymmetric bodies (namely, Sears-Haack body, semi-Sears-Haack body, 5-degree straight cone and flared cone) with different axial pressure gradients has been studied at Mach 2 in order to understand the effect of axial pressure gradient on instability amplification along the leeward symmetry plane and in the region of nonzero crossflow away from it. Comparisons are made with measured transition data in Mach 2 facilities as well as with predicted and measured transition characteristics for a 5-degree straight cone in a Mach 3.5 low disturbance tunnel. Limitations of using linear stability correlations for predicting transition over axisymmetric bodies at angle of attack are pointed out.
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Low-pressure blowoff experiments were conducted with a stagnation flame stabilized on the forward tip of cast PMMA rods in a vertical wind tunnel. Pressure, forced flow velocity, gravity, and ambient oxygen concentration were varied. Stagnation flame blowoff is determined from a time-stamped video recording of the test. The blowoff pressure is determined from test section pressure transducer data that is synchronized with the time stamp. The forced flow velocity is also determined from the choked flow orifice pressure. Most of the tests were performed in normal gravity, but a handful of microgravity tests were also conducted to determine the influence of buoyant flow velocity on the blowoff limits. The blowoff limits are found to have a linear dependence between the partial pressure of oxygen and the total pressure, regardless of forced flow velocity and gravity level. The flow velocity (forced and/or buoyant) affects the blowoff pressure through the critical Damkohler number residence time, which dictates the partial pressure of oxygen at blowoff. This is because the critical stretch rate increases linearly with increasing pressure at low pressure (sub-atmospheric pressures) since a second-order overall reaction rate with two-body reactions dominates in this pressure range.
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A high level overview of NASA’s X-59 Quesst mission will be presented as a lead in to invited special session on X-59 air data probe calibration efforts via wind tunnel testing, Neural Net development and Computation Fluid Dynamics.
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Explore the source record for details and available documents.
Explore the source record for details and available documents.
Explore the source record for details and available documents.
Nose cone structural and aerodynamic characteristics are essential for intelligent design of aircraft, spacecraft, and ballistic systems. Finite element analysis can be used to help understand the structural integrity and flight characteristics of different nose cones. A mesh sensitivity study was undertaken for a particular nose cone geometry that was used in tests at LANL facilities in order to confirm the integrity of the meshed geometry. A simple cone that best matched closed-form theoretical solutions was modeled, and received good correlation to the theory. Complexity was then added back to the nose cone. Parameters applied to the simple cone were then implemented in the nose cone geometry giving assurance of accuracy after the geometry was changed. Nose cone results averaged 6.3% error for radial displacement when compared with the theoretical. Hoop stress averaged 6.0% error and meridional stress averaged 5.7% error at the finest mesh level. Meshes showed signs of convergence when compared to all three theoretical solutions. Finally, pressure time-history data from LANL computational fluid dynamics simulations was applied to the surface of the final nose cone geometry. The pressure data was interpolated from pressure gauge locations onto nearby meshed elements, which allowed for FEA software to run simulations on the cone with the pressure data as a loading condition. The pressure mapping resulted in the ability to understand the nose cone’s rigid body motion that in turn can inform design of future nose cones.
Measurements of the local heat transfer and pressure distribution have been made on six 2-inch-diameter, blunt, axially symmetric bodies in the Langley gas dynamics laboratory at a Mach number of 4.95 and at Reynolds numbers per foot up to 81 x 10(exp 6). During the investigation laminar flow was observed over a hemispherical-nosed body having a surface finish from 10 to 20 microinches at the highest test Reynolds number per foot (for this configuration) of 77.4 x 10(exp 6). Though it was repeatedly possible to measure completely laminar flow at this Reynolds number for the hemisphere, it was not possible to observe completely laminar flow on the flat-nosed body for similar conditions. The significance of this phenomenon is obscured by the observation that the effects of particle impacts on the surface in causing roughness were more pronounced on the flat-nosed body. For engineering purposes, a method developed by M. Richard Dennison while employed by Lockheed Aircraft Corporation appears to be a reasonable procedure for estimating turbulent heat transfer provided transition occurs at a forward location on the body. For rearward-transition locations, the method is much poorer for the hemispherical nose than for the flat nose. The pressures measured on the hemisphere agreed very well with those of the modified Newtonian theory, whereas the pressures on all other bodies, except on the flat-nosed body, were bracketed by modified Newtonian theory both with and without centrifugal forces. For the hemisphere, the stagnation-point velocity gradient agreed very well with Newtonian theory. The stagnation-point velocity gradient for the flat- nosed model was 0.31 of the value for the hemispherical-nosed model. If a Newtonian type of flow is assumed, the ratio 0.31 will be independent of Much number and real-gas effects.
An investigation has been made in the Langley free-flight tunnel at low-subsonic speed to determine the static stability, control effectiveness, and damping in roll and yaw of a model with a low-aspect-ratio unswept wing and two different fuselage forebodies at angles of attack from 0 deg to 90 deg. Results were obtained with a fuselage configuration having a long pointed nose and a shorter rounded nose. Although the wing stalled at an angle of attack of about 12 deg, maximum lift did not occur until an angle of attack of about 40 deg or 50 deg was obtained. The static longitudinal stability of the model having a short rounded nose was greater than that of the model having a longer pointed nose over the entire angle-of-attack range. The pointed-nose model had large out-of-trim yawing moments above an angle of attack of about 40 deg. Shortening and rounding the nose of the model delayed these out-of-trim yawing moments to slightly higher angles of attack. Both models were directionally unstable above an angle of attack of about 20 deg, but both had positive effective dihedral over virtually the entire angle-of-attack range. At the higher angles of attack the pointed-nose model had generally better damping in roll than that of the rounded-nose model. Both models had very high damping in yaw at an angle of attack of about 50 deg or 60 deg.
The ability to measure occupants’ thermal state in real time will enable major advances in the control of air conditioning systems. This study proposes predicting occupant thermal state by a combination of infrared thermography, computer vision, and machine learning. The approach (1) uses cheek, nose, and hand temperatures because they are least subject to blockage by hair, glasses, and clothing; (2) measures the distribution of skin temperatures within geometrically defined sub-areas of the face and hand; and (3) uses temperature differences within and between these areas to eliminate the effects of calibration drift that are unavoidable in thermal infrared (TIR) cameras. Two series of tests were conducted, respectively in an outdoor carport and an indoor environmental chamber, collecting a total of 48,422 sets of cheek, nose, and hand skin temperatures using a TIR camera and computer-vision technology, coupled with 715 subjective responses of thermal sensations. To predict occupant thermal state, Random Forest classification models were built using either absolute skin temperatures (the maximum and median temperatures of cheek and hand segments, and the temperature of the central spot on the nose), or intra- and inter-segment temperature differences of cheeks, hands, and nose. These measurements were found to accurately predict occupant thermal state. Using the maximum and median temperatures for cheek and nose, or for cheek and hand, predicts thermal state with an accuracy of 92–96%. In conclusion, using only the intra- and inter-segment temperature differences from cheek and nose is 83% accurate; adding the hand temperature differences increases the accuracy to 96%.
Flight tests were performed on an F-14 aircraft to evaluate the use of flush pressure orifices on the nose section for obtaining air data at transonic speeds over a large range of flow angles. This program was part of a flight test and wind tunnel program to assess the accuracies of such systems for general use on aircraft. It also provided data to validate algorithms developed for the shuttle entry air data system designed at NASA Langley. Data were obtained for Mach numbers between 0.60 and 1.60, for angles of attack up to 26.0 deg, and for sideslip angles up to 11.0 deg. With careful calibration, a flush air data system with all flush orifices can provide accurate air data information over a large range of flow angles. Several orificies on the nose cap were found to be suitable for determination of stagnation pressure. Other orifices on the nose section aft of the nose cap were shown to be suitable for determination of static pressure. Pairs of orifices on the nose cap provided the most sensitive measurements for determining angles of attack and sideslip, although orifices located farther aft on the nose section could also be used.
Ice water content measurements were made simultaneously at three locations on the NASA DC-8 in natural, glaciated conditions during the 2018 High Ice Water Content RADAR flight campaign. The purpose of these measurements was to further evaluate efficiency factors of hot-wire total water content probes in glaciated conditions and investigate the enhancement of ice crystal concentrations near fuselage surfaces due to flow field inertial effects, and ice crystals impacting the nose, breaking up, and flowing downstream. The total water content measurements were made using either Science Engineering Associates Ice Crystal Detectors or Robust Probes. Three common sensors were mounted on an underwing canister considered to be in near free-flow conditions, a standoff from a fuselage window, and the nose of the fuselage near the pitot probes. The Ice Crystal Detector total water content sensor and Robust Probe sensor collection and retention efficiencies were evaluated through comparisons with the underwing Ice Crystal Detector and Robust Probe measurements to the IKP2 isokinetic evaporator probe, which provided the reference ice water content measurement. Local ice water content at the nose position was evaluated by comparing ratios of the nose and underwing ice crystal detectors to the IKP2. Local ice water content at the window-standoff location was also evaluated by comparing total water content sensor measurements from the window ice crystal detector to the measurements made with the underwing and nose Ice Crystal Detectors The key findings were: 1) the Ice Crystal Detector total water content sensor ice water content efficiency factor was similar to previous estimates, but reduced with increased ice crystal median mass diameter; 2) the ice water content at the fuselage nose location near the DC-8 pitot probes was approximately 2.5 times the freestream values – although this estimate is affected by a higher probe efficiency factor due to smaller particles in the debris cloud from impacts upstream of the probe; and 3) the ice water content at the 17” standoff from the port window varied from about 50% to 3 times freestream values in a complicated manner. Similar measurement locations are not uncommon on cloud research aircraft, where ice particle measurements may be subject to similar uncertainties.