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At least 235 records · Page 13

All sources NASA NTRS DOE OSTI

NASA NTRS1942

A method for determining the camber and twist of a surface to support a given distribution of lift

A graphical method is described for finding the shape (camber and twist) of an airfoil having an arbitrary distribution of lift. The method consists in replacing the lifting surface and its wake with an equivalent arrangement of vortices and in finding the associated vertical velocities. By a division of the vortex pattern into circular strips concentric about the downwash point instead of into the usual rectangular strips, the lifting surface is reduced for each downwash point to an equivalent loaded line for which the induced velocity is readily computed. The ratio of the vertical velocity to the stream velocity is the slope of the surface in the free-stream direction. As an illustration, the shape of the wing consistent with the pressure distribution derived from the two-dimensional theories is found for two wings: a straight elliptical wing and one with 30 degree sweepback. Application of the method to solve the reverse problem - finding the lift distribution over a given surface - is briefly discussed.

Cohen, Doris↗

NASA NTRS1993

Leading-edge transition and relaminarization phenomena on a subsonic high-lift system

Boundary-layer transition and relaminarization may have a critical effect on the flow development about multielement high-lift systems of subsonic transport aircraft with swept wings. The purpose of this paper is a study of transition phenomena in the leading-edge region of the various elements of a high-lift system. The flow phenomena studied include transition of the attachment-line flow, relaminarization, and crossflow instability. The calculations are based on pressure distributions measured in flight on the NASA Transport Systems Research Vehicle (Boeing 737-100) at a wing station where the flow approximated infinite swept wing conditions. The results indicate that significant regions of laminar flow can exist on all flap elements in flight. In future flight experiments the extent of these regions will be measured, and the transition mechanisms and the effect of laminar flow on the high-lift characteristics of the multi-element system will be further explored.

Van Dam, C. P.↗

NASA NTRS1993

Prediction of leading-edge transition and relaminarization phenomena on a subsonic multi-element high-lift system

Boundary-layer transition and relaminarization may have a critical effect on the flow development about multi-element high-lift systems of subsonic transport jets with swept wings. The purpose of the research is to study these transition phenomena in the leading-edge region of the various elements of a high-lift system. The flow phenomena studied include transition to the attachment-line flow, relaminarization, and crossflow instability, and transition. The calculations are based on pressure distributions measured in flight on the NASA Transport Systems Research Vehicle (Boeing 737-100) at a wing station where the flow approximated infinite swept wing conditions. The results indicate that significant regions of laminar flow can exist on all flap elements in flight. In future flight experiments (planned for January-February, 1994) the extent of these regions, the transition mechanisms and the effect of laminar flow on the high-lift characteristics of the multi-element system will be further explored.

Vandam, C. P.↗

NASA NTRS2002

Virtual Shaping of a Two-dimensional NACA 0015 Airfoil Using Synthetic Jet Actuator

The Aircraft Morphing Program at NASA Langley envisions an aircraft without conventional control surfaces. Instead of moving control surfaces, the vehicle control systems may be implemented with a combination of propulsive forces, micro surface effectors, and fluidic devices dynamically operated by an intelligent flight control system to provide aircraft maneuverability over each mission segment. As a part of this program, a two-dimensional NACA 0015 airfoil model was designed to test mild maneuvering capability of synthetic jets in a subsonic wind tunnel. The objective of the experiments is to assess the applicability of using unsteady suction and blowing to alter the aerodynamic shape of an airfoil with a purpose to enhance lift and/or to reduce drag. Synthetic jet actuation at different chordwise locations, different forcing frequencies and amplitudes, under different freestream velocities are investigated. The effect of virtual shape change is indicated by a localized increase of surface pressure in the neighborhood of synthetic jet actuation. That causes a negative lift to the airfoil with an upper surface actuation. When actuation is applied near the airfoil leading edge, it appears that the stagnation line is shifted inducing an effect similar to that caused by a small angle of attack to produce an overall lift change.

Chen, Fang-Jenq↗

NASA NTRS1977

Unsteady supersonic aerodynamic theory by the method of potential gradient

A generalized solution of the hyperbolic wave equation has been derived. The method used has been further developed to relate the velocity components at a field point to the potential gradient distribution in the dependence domain. Singular integrals have been evaluated in closed form, while numerical integration methods are suggested for treating more complex but analytic functions. Idealization of the lifting surfaces by trapezoidal elements with two sides parallel to the streamlines is computationally efficient because line integrations along the other two sides need only be considered. Furthermore, all the integrands vanish on the Mach cone and the need for determining the hyperbolic curves of intersection of the cone with the lifting surface is avoided. Generalized aerodynamic coefficients for three AGARD planforms have been calculated and compared with the available results.

Jones, W. P.↗

NASA NTRS1957

Aerodynamic Forces and Moments on a Large Ogive-Cylinder Store at Various Locations Below the Fuselage Center Line of a Swept-Wing Bomber Configuration at a Mach Number of 1.61

A supersonic wind-tunnel investigation on store interference has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.61. Forces and moments were measured on a large ogive-cylinder store in the presence of a 45 deg swept-wing-fuselage bomber configuration for a number of store locations below the fuselage center line. Results of the investigation show that large variations of store lift, drag, and pitch occur with changes in store or airplane angle of attack, store vertical location, and store horizontal location. The variation of the store forces and moments with respect to the chordwise location of the wing plan form indicates that the wing is a large factor in producing the interference loads on the store. Comparison of data for underfuselage and underwing store locations at an angle of attack of 0 deg showed maximum store drag interferences of similar magnitudes, but showed considerably smaller maximum interference on store lift an pitching moments for underfuselage store locations.

SUPERSONIC FLIGHT↗

NASA NTRS1987

Calculating Sonic-Boom Propagation

Nonlinear effects included, enabling more-realistic modeling. Modified Method of Characteristics Sonic Boom Extrapolation Program (MMOC) is computer program for sonic-boom propagation that includes shock coalescence and incorporates effects of asymmetries due to volume and lift. Numerically integrates nonlinear governing equations using data on initial data line approximately one body length from aircraft and yields sonic-boom pressure at ground as function of time or of position at given time. MMOC written in FORTRAN IV.

Darden, C. M.↗

NASA NTRS2020

An assessment of Earth's climate sensitivity using multiple lines of evidence

We assess evidence relevant to Earth's equilibrium climate sensitivity per doubling of atmospheric CO2, characterized by an effective sensitivity S . This evidence includes feedback process understanding, the historical climate record, and the paleoclimate record. An S value lower than 2 K is difficult to reconcile with any of the three lines of evidence. The amount of cooling during the Last Glacial Maximum provides strong evidence against values of S greater than 4.5 K. Other lines of evidence in combination also show that this is relatively unlikely. We use a Bayesian approach to produce a probability density (PDF) for S given all the evidence, including tests of robustness to difficult‐to‐quantify uncertainties and different priors. The 66% range is 2.6‐3.9 K for our Baseline calculation, and remains within 2.3‐4.5 K under the robustness tests; corresponding 5‐95% ranges are 2.3‐4.7 K, bounded by 2.0‐5.7 K (although such high‐confidence ranges should be regarded more cautiously). This indicates a stronger constraint on S than reported in past assessments, by lifting the low end of the range. This narrowing occurs because the three lines of evidence agree and are judged to be largely independent, and because of greater confidence in understanding feedback processes and in combining evidence. We identify promising avenues for further narrowing the range in S , in particular using comprehensive models and process understanding to address limitations in the traditional forcing‐feedback paradigm for interpreting past changes.

Climate↗

DOE OSTI2020

An assessment of Earth's climate sensitivity using multiple lines of evidence

In this work, we assess evidence relevant to Earth's equilibrium climate sensitivity per doubling of atmospheric CO 2 , characterized by an effective sensitivity S . This evidence includes feedback process understanding, the historical climate record, and the paleoclimate record. An S value lower than 2 K is difficult to reconcile with any of the three lines of evidence. The amount of cooling during the Last Glacial Maximum provides strong evidence against values of S greater than 4.5 K. Other lines of evidence in combination also show that this is relatively unlikely. We use a Bayesian approach to produce a probability density (PDF) for S given all the evidence, including tests of robustness to difficult-to-quantify uncertainties and different priors. The 66% range is 2.6-3.9 K for our Baseline calculation, and remains within 2.3-4.5 K under the robustness tests; corresponding 5-95% ranges are 2.3-4.7 K, bounded by 2.0-5.7 K (although such high-confidence ranges should be regarded more cautiously). This indicates a stronger constraint on S than reported in past assessments, by lifting the low end of the range. This narrowing occurs because the three lines of evidence agree and are judged to be largely independent, and because of greater confidence in understanding feedback processes and in combining evidence. We identify promising avenues for further narrowing the range in S , in particular using comprehensive models and process understanding to address limitations in the traditional forcing-feedback paradigm for interpreting past changes.

54 ENVIRONMENTAL SCIENCES↗

NASA NTRS1958

Lift, Drag, and Pitching Moment of an Aspect-Ratio-2 Triangular Wing with Leading-Edge Flaps Designed to Simulate Conical Camber

An investigation was conducted to determine the effectiveness of leading-edge flaps in reducing the drag at lifting conditions of a triangular wing of aspect ratio 2.0. The flaps, deflected to simulate conically cambered wings having a wide range of design lift coefficients, were tested over a Mach number range of 0.70 to 2.22 through an angle-of-attack variation from -6 deg to +18 deg at a constant Reynolds number of 3.68 million based on the wing mean aerodynamic chord. A symmetrical wing of the same plan form and aspect ratio was also tested to provide a basis for comparison. The experimental results showed that with the flaps in the undeflected position, a small amount of fixed leading-edge droop incorporated over the outboard 5 percent of the wing semispan was as effective at high subsonic speeds as conical camber in improving the maximum lift-drag ratio above that of the symmetrical wing. At supersonic speeds, the penalty in minimum drag above that of the symmetrical wing was less than that incurred by conical camber. Deflecting the leading-edge flaps about the hinge line through 80 percent of the wing semispan resulted in further improvements of the drag characteristics at lift coefficients above 0.20 throughout the Mach number range investigated. The lift and pitching-moment characteristics were not significantly affected by the leading-edge flaps.

Menees, Gene P.↗

NASA NTRS1975

Integrated potential formulation of unsteady supersonic aerodynamics for interacting wings

A numerical integrated velocity potential method for the determination of unsteady aerodynamic forces on arbitrary interacting wings and tails in supersonic flow has been developed. Constant Mach number has been assumed throughout the flow field. Normalwash and sidewash integrals have been derived. The upwash integral remains to be derived. Singular integrals in the expressions for the velocity components have been evaluated in closed form. Lifting surfaces are represented by triangular elements defined by arbitrarily spaced characteristic lines and the true surface edges. The wake field is represented by rectangular strip elements. Velocity potential distributions and generalized aerodynamic coefficients have been compared.

Appa, K.↗

NASA NTRS1979

Comparison of three-dimensional panel methods with strip boundary-layer simulations to experiment

Several three dimensional surface panel methods coupled on a streamwise strip basis to two dimensional boundary layer methods were compared to experimental data. Several different boundary layer simulations were investigated. Comparable boundary layer effects which converge to the expected two dimensional result wind increasing aspect ratio were indicated by either a transpiration or camber line displacement simulation. The method was applied to several varied configurations. The lift predictions agreed well with an approximate method based on correcting the inviscid lift using the two dimensional section viscous characteristics. The wing pressures and the effect of the boundary layer on the wing pressures were predicted accurately.

Kjelgaard, S. O.↗

NASA NTRS1983

Rolling Beam Umbilical System

The decision to make the Centaur Vehicle a Space Transportation System (STS) payload meant that new ground support equipment provisions at Launch Pads Pads 39A and 39B were required. These new equipment provisions were needed to service the Centaur vehicle while it was installed in the Orbiter's payload bay prior to launch. This paper describes the design of a new rolling beam umbilical syste (RBUS) being added to the pad fixed service structure (FSS) in order to provide the primary functions of liquid hydrogen (LH2) fill, drain, and vent. The carrier plate itself is a Government-furnished equipment item and of necessity became a T-O disconnect. This permits quick offloading in the event of an abort prior to lift-off. In addition to the rolling beam structure, mechanisms, and fluid lines, it was necessary to design and build a carrier plate simulator to support early development testing of the mast at the Launch Equipment Test Facility at Kennedy Space Center. The RBUS is designed to be compatible with the rotating service structure (RSS) to the extent that the umbilical may be deployed with the RSS mated with the vehicle. It is also designed to clear the RSS as the RSS rotates back out of the way. Accessibility to the Orbiter aft compartment via the 50-1 door had to be maintained.

Tatem, B. C., Jr.↗

NASA NTRS1987

Control of vortical separation on conical bodies

In a variety of aeronautical applications, the flow around conical bodies at incidence is of interest. Such applications include, but are not limited to, highly maneuverable aircraft with delta wings, the aerospace plane and nose portions of spike inlets. The theoretical model used has three parts. First, the single line vortex model is used within the framework of slender body theory, to compute the outer inviscid field for specified separation lines. Next, the three dimensional boundary layer is represented by a momentum equation for the cross flow, analogous to that for a plane boundary layer; a von Karman Pohlhausen approximation is applied to solve this equation. The cross flow separation for both laminar and turbulent layers is determined by matching the pressure at the upper and lower separation points. This iterative procedure yields a unique solution for the separation lines and consequently for the position of the vortices and the vortex lift on the body. Lastly, control of separation is achieved by blowing tangentially from a slot located along a cone generator. It is found that for very small blowing coefficients, the separation can be postponed or suppressedy completely.

Mourtos, Nikos J.↗

NASA NTRS1991

Airborne rescue system

The airborne rescue system includes a boom with telescoping members for extending a line and collar to a rescue victim. The boom extends beyond the tip of the helicopter rotor so that the victim may avoid the rotor downwash. The rescue line is played out and reeled in by winch. The line is temporarily retained under the boom. When the boom is extended, the rescue line passes through clips. When the victim dons the collar and the tension in the line reaches a predetermined level, the clips open and release the line from the boom. Then the rescue line can form a straight line between the victim and the winch, and the victim can be lifted to the helicopter. A translator is utilized to push out or pull in the telescoping members. The translator comprises a tape and a rope. Inside the telescoping members the tape is curled around the rope and the tape has a tube-like configuration. The tape and rope are provided from supply spools.

Haslim, Leonard A.↗

NASA NTRS1945

A method for determining the camber and twist of a surface to support a given distribution of lift, with applications to the load over a sweptback wing

Cohen, Doris↗

NASA NTRS1952

Theoretical symmetric span loading due to flap deflection for wings of arbitrary plan form at subsonic speeds

A simplified lifting-surface theory is applied to the problem of evaluating span loading due to flap deflection for arbitrary wing plan forms. With the resulting procedure, the effects of flap deflection on the span loading and associated aerodynamic characteristics can be easily computed for any wing which is symmetrical about the root chord and which has a straight quarter-chord line over the wing semispan. The effects of compressibility and spanwise variation of section lift-curve slope are taken into account by the procedure. The method presented can also be used to calculate the downwash in the vertical center of the wake of a wing which has arbitrary spanwise loading.

Deyoung, John↗

NASA NTRS1936

Correction of downwash in wind tunnels of circular and elliptic sections

The downwash velocity distribution behind the wing was determined for the free jet and for the closed tunnel of both circular and elliptic cross sections. The wing was placed at the center of the tunnel. The theory makes it possible to determine the downwash at any point in the jet. The computations were performed for points in the plane determined by the jet axis and the center-of-pressure line of the wing. The downwash proved to be proportional to the wing lift and inversely proportional to the cross-sectional area of the tunnel.

Lotz, Irmgard↗