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At least 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

Some Aspects of the Modified Newtonian and Prandtl-Meyer-Expansion Method for Axisymmetric Blunt Bodies at Zero Angle of Attack

In a recent article on hypersonic flow, Lees and Kubota suggested a modification of the local-expansion method to predict the surface pressures in the region of a nose-cylinder junction where Newtonian theory fails. The modification consists of matching the Newtonian pressure distribution and the Prandtl-Meyer relation at the point on the nose surface where both the pressure and the pressure gradient, predicted by these two solution, are equal. With this approximation, the pressure distribution can be predicted quite accurately up to the nose-cylinder junction for blunt-nosed bodies of continuous curvature, provided centrifugal-force effects are negligible. The purpose of this note is to facilitate the application of Lees' suggestion and to draw attention to some of its consequences.

Wagner, Richard D., Jr.↗
NASA NTRS2016

Characterization of Carbon Dioxide Washout Measurement Techniques in the Mark-III Space Suit

A space suit must provide adequate carbon dioxide (CO2) washout inside the helmet to prevent symptoms of hypercapnia. In the past, an oronasal mask has been used to measure the inspired air of suited subjects to determine a space suit's CO2 washout capability. While sufficient for super-ambient pressure testing of space suits, the oronasal mask fails to meet several human factors and operational criterion needed for future sub-ambient pressure testing (e.g. compatibility with a Valsalva device). This paper describes the evaluation of a nasal cannula as a device for measuring inspired air within a space suit. Eight test subjects were tasked with walking on a treadmill or operating an arm ergometer to achieve target metabolic rates of 1000, 2000, and 3000 British thermal units per hour (BTU/hr), at flow rates of 2, 4, and 6 actual cubic feet per minute (ACFM). Each test configuration was conducted twice, with subjects instructed to breathe either through their nose only, or however they felt comfortable. Test data shows that the nasal cannula provides more statistically consistent data across test subjects than the oronasal mask used in previous tests. The data also shows that inhaling/exhaling through only the nose provides a lower sample variance than a normal breathing style. Nose-only breathing reports better CO2 washout due to several possible reasons, including a decreased respiratory rate, an increased tidal volume, and because nose-only breathing directs all of the exhaled CO2 down and away from the oronasal region. The test subjects in this study provided feedback that the nasal cannula is comfortable and can be used with the Valsalva device.

Meginnis, Ian M.↗
NASA NTRS2016

Crash Testing and Simulation of a Cessna 172 Aircraft: Hard Landing Onto Concrete

A full-scale crash test of a Cessna 172 aircraft was conducted at the Landing and Impact Research Facility at NASA Langley Research Center during the summer of 2015. The purpose of the test was to evaluate the performance of Emergency Locator Transmitters (ELTs) that were mounted at various locations in the aircraft and to generate impact test data for model validation. A finite element model of the aircraft was developed for execution in LSDYNA to simulate the test. Measured impact conditions were 722.4-in/s forward velocity and 276-in/s vertical velocity with a 1.5deg pitch (nose up) attitude. These conditions were intended to represent a survivable hard landing. The impact surface was concrete. During the test, the nose gear tire impacted the concrete, followed closely by impact of the main gear tires. The main landing gear spread outward, as the nose gear stroked vertically. The only fuselage contact with the impact surface was a slight impact of the rearmost portion of the lower tail. Thus, capturing the behavior of the nose and main landing gear was essential to accurately predict the response. This paper describes the model development and presents test-analysis comparisons in three categories: inertial properties, time sequence of events, and acceleration and velocity time-histories.

Jackson, Karen E.↗
NASA NTRS2019

Nonmodal Growth of TravelingWaves on Blunt Cones at Hypersonic Speeds

The existing database of transition measurements in hypersonic ground facilities has established that, as the nosetip bluntness is increased, the onset of boundary layer transition over a circular cone at zero angle of attack shifts downstream. However, this trend is reversed at sufficiently large values of the nose Reynolds number, so that the transition onset location eventually moves upstream with a further increase in nose-tip bluntness. Because modal amplification is too weak to initiate transition at moderate-to-large bluntness values, nonmodal growth has been investigated as the potential basis for a physics-based model for the frustum transition. The present analysis investigates the nonmodal growth of traveling disturbances initiated within the nose-tip vicinity that peak within the entropy layer. Results show that, with increasing nose bluntness, both planar and oblique traveling disturbances experience appreciable energy amplification up to successively higher frequencies. For moderately blunt cones, the initial nonmmodal growth is followed by a partial decay that is more than overcome by an eventual, modal growth as Mack-mode waves. For larger bluntness values, the Mack-mode waves are not amplified anywhere upstream of the experimentally measured transition location, but the traveling modes still undergo a significant amount of nonmodal growth. This finding does not provide a definitive link between optimal growth and the onset of transition, but it is qualitatively consistent with the experimental observations that frustum transition in the absence of sufficient Mack-mode amplification implies a double peak in disturbance amplification and the appearance of transitional events above the boundary-layer edge.

Hypersonic speed↗
NASA NTRSDate not supplied

Tracking the Subsolar Bow Shock and Magnetopause

Global magnetohydrodynamic models predict that plasma velocities vary almost linearly from 0 km s-1 along the stagnation streamline leading from the nose of the stationary magnetopause to 0.25 Vsw at the nose of the bow shock, where Vsw is the solar wind velocity. This paper presents a proof-of-concept study showing how two-point measurements of the plasma velocity in the subsolar magnetosheath can be used to determine gradients in the plasma velocity and consequently the steady-state or slowly-moving locations of both the nose of the magnetopause and the nose of the bow shock. The results may be of use to those binning magnetosheath observations to develop empirical models, those testing models for magnetopause erosion or deposition in response to magnetic reconnection, and those determining the stand-off distance of the bow shock as a function of solar wind conditions.

David Sibeck↗
NASA NTRS1951

Determination of Shapes of Boattail Bodies of Revolution for Minimum Wave Drag

By use of an approximate equation for the wave drag of slender bodies of revolution in a supersonic flow field, the optimum shapes of certain boattail bodies are determined for minimum wave drag. The properties of three specific families of bodies are determined, the first family consisting of bodies having a given length and base area and a contour passing through a prescribed point between the nose and base, the second family having fixed length, base area, and maximum area, and the third family having given length, volume, and base area. The method presented is easily generalized to determine minimum-wave-drag profile shapes which have contours that must pass through any prescribed number of points. According to linearized theory, the optimum profiles are found to have infinite slope at the nose but zero radius of curvature so that the bodies appear to have pointed noses, a zero slope at the body base, and no variation of wave drag with Mach number. For those bodies having a specified intermediate.diameter (that is, location and magnitude given), the maximum body diameter is shown to be larger, in general, than the specified diameter. It is also shown that, for bodies having a specified maximum diameter, the location of the maximum diameter is not arbitrary but is determined from the ratio of base diameter to maximum diameter.

Adams, Mac C.↗
NASA NTRS1972

Asymmetric lateral-directional characteristics of pointed bodies of revolution at high angles of attack

A low-speed wind-tunnel investigation was conducted in order to determine the cause of asymmetric yawing moments produced by long pointed fuselage nose shapes at high angles of attack. Force tests were conducted with a cone, tangent-ogive model, and a paraboloid of revolution over a range of Reynolds numbers from 0.15 million to 0.35 million for an angle-of-attack range from 0 deg to 75 deg and an angle-of-sideslip range of plus or minus 30 deg. Tuft and smoke flow-visualization tests were also conducted to aid in the analysis. Large asymmetric yawing moments were obtained for the cone and tangent-ogive body at high angles of attack (of the order of 40 to 60 deg). These large moments were caused by asymmetric shedding of vortex sheets off the long pointed nose. The asymmetric moments could be eliminated by use of symmetrically arranged strikes on the nose. The paraboloid of revolution did not produce a strong asymmetric flow field at high angles of attack and did not exhibit asymmetric moments.

Coe, P. L., Jr.↗
NASA NTRS1973

Local skin friction coefficients and boundary layer profiles obtained in flight from the XB-70-1 airplane at Mach numbers up to 2.5

Boundary-layer and local friction data for Mach numbers up to 2.5 and Reynolds numbers up to 3.6 x 10 to the 8th power were obtained in flight at three locations on the XB-70-1 airplane: the lower forward fuselage centerline (nose), the upper rear fuselage centerline, and the upper surface of the right wing. Local skin friction coefficients were derived at each location by using (1) a skin friction force balance, (2) a Preston probe, and (3) an adaptation of Clauser's method which derives skin friction from the rake velocity profile. These three techniques provided consistent results that agreed well with the von Karman-Schoenherr relationship for flow conditions that are quasi-two-dimensional. At the lower angles of attack, the nose-boom and flow-direction vanes are believed to have caused the momentum thickness at the nose to be larger than at the higher angles of attack. The boundary-layer data and local skin friction coefficients are tabulated. The wind-tunnel-model surface-pressure distribution ahead of the three locations and the flight surface-pressure distribution ahead of the wing location are included.

Fisher, D. F.↗
NASA NTRS1973

Overall fluctuating pressure levels on prospective space shuttle launch configurations at Mach numbers from 0.8 to 2.2

Overall fluctuating pressure levels of seven space shuttle launch configurations are presented. The model was a 4-percent-scale space shuttle vehicle, tested in both a 11- by 11-foot transonic wind tunnel and a 9- by 7-foot supersonic wind tunnel. Mach numbers varied from 0.8 to 2.2, and the angle of attack range was from -8 deg to 8 deg at angles of sideslip of -5 deg, and 5 deg. The model configurations included both series-burn and parallel-burn configurations, two canopy configurations, two positions of the orbiter nose relative to the HO tank nose and two HO tank nose-cone angles (15 deg and 20 deg). The fluctuating pressure levels are presented in three forms.

Dods, J. B., Jr.↗
NASA NTRS1974

Wind-tunnel free-flight investigation of a model of a spin-resistant fighter configuration

An investigation was conducted to provide some insight into the features affecting the high-angle-of-attack characteristics of a high-performance twin-engine fighter airplane which in operation has exhibited excellent stall characteristics with a general resistance to spinning. Various techniques employed in the study included wind-tunnel free-flight tests, flow-visualization tests, static force tests, and dynamic (forced-oscillation) tests. In addition to tests conducted on the basic configuration tests were made with the wing planform and the fuselage nose modified. The results of the study showed that the model exhibited good dynamic stability characteristics at angles of attack well beyond that for wing stall. The directional stability of the model was provided by the vertical tail at low and moderate angles of attack and by the fuselage forebody at high angles of attack. The wing planform was found to have little effect on the stability characteristics at high angles of attack. The tests also showed that although the fuselage forebody produced beneficial contributions to static directional stability at high angles of attack, it also produced unstable values of damping in yaw. Nose strakes located in a position which eliminated the beneficial nose contributions produced a severe directional divergence.

Grafton, S. B.↗
NASA NTRS1975

Experimental aerodynamics characteristics for bodies of elliptic cross section at angles of attack from 0 deg to 58 deg and Mach numbers from 0.6 to 2.0

An experimental investigation was conducted to measure the static aerodynamic characteristics for two bodies of elliptic cross section and for their equivalent body of revolution. The equivalent body of revolution had the same length and axial distribution of cross-sectional area as the elliptic bodies. It consisted of a tangent ogive nose of fineness ratio 3 followed by a cylinder with a fineness ratio of 7. All bodies were tested at Mach numbers of 0.6, 0.9, 1.2, 1.5, and 2.0 at angles of attack from 0 deg to 58 deg. The data demonstrate that the aerodynamic characteristics can be significantly altered by changing the body cross section from circular to elliptic and by rolling the body from 0 deg to 90 deg. For example, the first elliptic body (with a constant cross-sectional axis ratio of 2) developed at zero roll about twice the normal force developed by the equivalent body of revolution. At some angles of attack greater than about 25 deg, side forces and yawing moments were measured in spite of the fact that the bodies were tested at zero angle of sideslip. The side-force and yawing-moment coefficients decreased with an increase in Mach number and essentially disappeared for all the bodies at Mach numbers greater than 1.2. From the standpoint of reducing undesirable side forces at high angles of attack, it is best to have the flattest side of the nose of the elliptic bodies pitching against the stream crossflow. The effect of Reynolds number was also the least significant for both elliptic bodies when the flattest side of the nose was pitched against the stream crossflow.

Jorgensen, L. H.↗
NASA NTRS1975

Investigation of space shuttle vehicle 140C configuration orbiter (model 16-0) wheel well pressure loads in the Rockwell International 7.75 x 11 foot wind tunnel (OA143)

Experimental aerodynamic investigations were conducted on a sting mounted .0405-scale representation of the 140C outer mold line space shuttle orbiter configuration in the Rockwell International 7.75 x 11.00 foot low speed wind tunnel. The primary test objectives were to define the orbiter wheel well pressure loading and its effects on landing gear thermal insulation and to investigate the pressure environment experienced by both the horizontal flight nose probe and air vent door probes. Steady state and dynamic pressure values were recorded in the orbiter nose gear well, left main landing gear well, horizontal flight nose probe, and both left and right air vent door probe. All steady state pressure levels were measured by Statham differential pressure transducers while dynamic pressure levels were recorded by Kulite high frequency response pressure sensors.

Mennell, R. C.↗
NASA NTRS1975

Shadowgraphs of air flow over prospective space shuttle configurations at Mach numbers from 0.8 to 1.4

Shadowgraphs of five space shuttle launch configurations are presented. The model was a 4 percent-scale space shuttle vehicle, tested in the 11- by 11-foot Transonic Wind Tunnel at Ames Research Center. The Mach number was varied from 0.8 to 1.4 with three angles of sideslip (0 deg, 5 deg and -5 deg) that were used in conjunction with three angles of attack (4 deg, -4 deg, and 0 deg). The model configurations included both series-burn and parallel-burn configurations, two canopy configurations, two positions of the orbiter nose relative to the HO tank nose, and two HO tank nose-cones angles (15 deg and 20 deg). The data consist entirely of shadowgraph photographs.

Dods, J. B., Jr.↗
NASA NTRS1975

Laminar separation on a blunted cone at high angles of attack

A new technique is reported for calculating the entire flow field on spherically blunted cones at high angles of attack and high laminar Reynolds numbers. An approximate system of parabolic equations obtained from the steady Navier-Stokes equations by assuming the viscous, streamwise derivative terms are small compared to the viscous normal and circumferential derivatives is the basis of the calculations. These equations are valid for both the inviscid and viscous regions, including the circumferential separation zone that develops on the leeward side at high angles of attack. Two different methods are used to obtain the initial conditions for these equations at the sphere cone tangency plane. For small nose Reynolds numbers, an axisymmetric merged layer solution around a sphere is rotated to provide a three-dimensional initial plane of data. For large nose Reynolds numbers, the nose region is solved using an inviscid, three dimensional time dependent solution combined with a boundary layer solution for the viscous flow. The computed flowfield including the leeward separation region is described and compared with data for a 7 deg half angle cone at 10 deg angle of attack, and a blunt 15 deg half angle cone at 15 deg angle of attack.

Lubard, S. C.↗
NASA NTRS1975

Wind tunnel investigation of the aerodynamic characteristics of five forebody models at high angles of attack at Mach numbers from 0.25 to 2

Five forebody models of various shapes were tested in the Ames 6- by 6-Foot Wind Tunnel to determine the aerodynamic characteristics at Mach numbers from 0.25 to 2 at a Reynolds number of 800000. At a Mach number of 0.6 the Reynolds number was varied from 0.4 to 1.8 mil. Angle of attack was varied from -2 deg to 88 deg at zero sideslip. The purpose of the investigation was to determine the effect of Mach number of the side force that develops at low speeds and zero sideslip for all of these forebody models when the nose is pointed. Test results show that with increasing Mach number the maximum side forces decrease to zero between Mach numbers of 0.8 and 1.5, depending on the nose angle; the smaller the nose angle of the higher the Mach number at which the side force exists. At a Mach number of 0.6 there is some variation of side force with Reynolds number, the variation being the largest for the more slender tangent ogive.

Keener, E. R.↗
NASA NTRS1975

An evaluation of the MSBLS/nose boom interference for the approach and landing test

An analysis was performed to investigate whether or not the obstruction caused by boom-mounted probes forward of the Orbiter nose could cause Microwave Scan Beam Landing System (MSBLS) data dropout during the autoland flight phase of the Approach and Landing Test (ALT). The results indicate that there will be no interference from the nose boom to cause MSBLS data dropout for the flight conditions under which the ALT is to be performed. Additional analysis considering approach trajectories in more extreme wind conditions than anticipated for the ALT also indicates that the nose boom creates no interference.

Tate, J. M.↗
NASA NTRS1976

Experimental investigation of oscillations in flows over shallow cavities

Wind tunnel experiments were carried out with three axisymmetric models to study laminar axisymmetric flows over shallow cavities at low subsonic speeds. The first model had a hemispherical nose, the second an ogive-shape nose, and the third an elliptic nose. Constant-temperature hot-wire anemometry was used in measuring both mean and fluctuating quantities. Flow around the cavity was visualized by heating the first model. To study the effect of mass injection on cavity oscillations, air was injected circumferentially all along the base of the cavity. Major results are that the cavity depth has little effect on oscillations in shallow cavities except when the depth is of the order of the thickness of the cavity shear flow, that downstream corner is the key factor in inducing self-sustained oscillations in the cavity shear layer, and that the presence of a back face results in an integral relation between wave length of the propagating disturbances and the cavity width in each mode of cavity operation.

Sarohia, V.↗
NASA NTRS1977

Side forces on a tangent ogive forebody with a fineness ratio of 3.5 at high angles of attack and Mach numbers from 0.1 to 0.7

An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics, at high angles of attack, of a tangent ogive forebody with a fineness ratio of 3.5. The investigation included the effects of nose bluntness, nose strakes, nose booms, a simulated canopy, and boundary-layer trips. The forebody was also tested with a short afterbody attached. Static longitudinal and lateral-directional stability data were obtained at Reynolds numbers ranging from 0.3 mil. to 3.8 mil. (based on base diameter) at a Mach number of 0.25, and at a Reynolds number of 0.8 mil. at Mach numbers ranging from 0.1 to 0.7. Angle of attack was varied from 0 to 88 deg at zero sideslip, and the sideslip angle was varied from -10 to 30 deg at angles of attack of 40, 55, and 70 deg.

Keener, E. R.↗