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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Computational study of a lunar return aerobrake concept

The flow field around a proposed lunar return aerobrake is examined computationally assuming viscous, laminar flow and utilizing an effective gamma approach to incorporate real-gas effects. The flow fields for three cases are calculated with both an axisymmetric and a three-dimensional formulation. The three vehicle configurations have braking panels extended 50 deg, 55 deg, and 60 deg to the inflow (zero lift configurations). The resulting axisymmetric and 3D flow fields are examined, and it is shown that despite complexities in the 3D flow field, the total drag coefficients calculated from modified axisymmetric results are very close to those obtained from the 3D solutions. The aerobrake is shown to achieve total drag coefficients as high as 8.4 for a vehicle with panels deflected to 60 deg.

Polsky, Susan↗
NASA NTRS1974

Transonic conical flow

The problem of inviscid, steady transonic conical flow, formulated in terms of the small disturbance theory, is studied. The small disturbance equation and similarity rules are presented, and a boundary value problem is formulated for the case of a supersonic freestream Mach number. The equation for the perturbation potential is solved numerically using an elliptic finite difference system. The difference equations are solved with a point relaxation algorithm that is also capable of capturing the shock wave during the iteration procedure by using the boundary conditions at the shock. Numerical calculations, for shock location, pressure distribution and drag coefficient, are presented for a family of nonlifting conical wings. The theory of slender wings is also presented and analytical results for pressure and drag coefficients are obtained.

Agopian, K. G.↗
NASA NTRS1975

Reduction of trimmed drag

Methods are reported for reducing the aircraft drag coefficient for a given aircraft lift coefficient, or speed. The emphasis is placed in determining the load distribution between the wing-body combination and the tail which reduces overall drag coefficient. Furthermore, a technique is presented which allows the determination of various aerodynamic and geometric parameters to permit the best location to satisfy inherent stability requirements. Included in the method is the calculation of sensitivity coefficients which indicate the importance of various parameters in achieving specified goals. Preliminary results indicate that such an approach is feasible.

Lutze, F. H., Jr.↗
NASA NTRS1985

Icing flight research: Aerodynamic effects of ice and ice shape documentation with stereo photography

Aircraft icing flight research was performed in natural icing conditions. A data base consisting of icing cloud measurements, ice shapes, and aerodynamic measurements is being developed. During research icing encounters the icing cloud was continuously measured. After the encounter, the ice accretion shapes on the wing were documented with a stereo camera system. The increase in wing section drag was measured with a wake survey probe. The overall aircraft performance loss in terms of lift and drag coefficient changes was obtained by steady level speed/power measurements. Selective deicing of the airframe components was performed to determine their contributions to the total drag increase. Engine out capability in terms of power available was analyzed for the iced aircraft. It was shown that the stereo photography system can be used to document ice shapes in flight and that the wake survey probe can measure increases in wing section drag caused by ice. On one flight, the wing section drag coefficient (c sub d) increased approximately 120 percent over the uniced baseline at an aircraft angle of attack of 6 deg. On another flight, the aircraft darg coefficient (c sub d) increased by 75 percent over the uniced baseline at an aircraft lift coefficient (C sub d) of 0.5.

Mikkelsen, K. L.↗
NASA NTRS1992

CFD modeling of turbulent flows around the SSME main injector assembly using porosity formulation

Hot gas turbulent flow distribution around the main injector assembly of the Space Shuttle Main Engine (SSME) and LOX flow distribution through the LOX posts have a great effect on the combustion phenomena inside the main combustion chamber. In order to design a CFD model to be an effective engineering analysis tool with good computational turn-around time (especially for 3-D flow problems) and still maintain good accuracy in describing the flow features, the concept of porosity was employed to describe the effects of blockage and drag force due to the presence of the LOX posts in the turbulent flow field around the main injector assembly of the SSME. Two-dimensional numerical studies were conducted to identify the drag coefficients of the flows, both through tube banks and round the shielded posts, over a wide range of Reynolds numbers. Empirical, analytical expressions of the drag coefficients as a function of local flow Reynolds number were then deduced. The porosity model was applied to the turbulent flow around the main injector assembly of the SSME, and analyses were performed. The 3-D CFD analysis was divided into three parts: LOX dome, hot gas injector assembly, and hydrogen cavity. The numerical results indicate that the mixture ratio at the downstream of injector face was close to stoichiometric around baffle elements.

Cheng, Gary C.↗
NASA NTRS1995

Flow Distribution Around the SSME Main Injector Assembly Using Porosity Formulation

Hot gas turbulent flow distribution around the main injector assembly of the Space Shuttle Main Engine (SSME) and Liquid Oxidizer (LOX) flow distribution through the LOX posts have a great effect on the combustion phenomena inside the main combustion chamber. In order to design a CFD model to be an effective engineering analysis tool with good computational turn-around time (especially for 3-D flow problems) and still maintain good accuracy in describing the flow features, the concept of porosity was employed to describe the effects of blockage and drag force due to the presence of the LOX posts in the turbulent flow field around the main injector assembly of the SSME. 2-D numerical studies were conducted to identify the drag coefficients of the flows both through tube banks and around the shielded posts over a wide range of Reynolds numbers. Empirical, analytical expressions of the drag coefficient as a function of local flow Reynolds number were then deduced. The porosity model was applied to the turbulent flow around the main injector assembly of the SSME, and analyses were performed. The 3-D CFD analysis was divided into three parts, LOX dome, hot gas injector assembly, and hydrogen cavity. The numerical results indicate that the mixture ratio at the downstream of injector face was close to stoichiometric around baffle elements.

Cheng, Gary C.↗
NASA NTRS1995

Flow Distribution Around the SSME Main Injector Assembly Using Porosity Formulation

Hot gas turbulent flow distribution around the main injector assembly of the Space Shuttle Main Engine (SSME) and Liquid Oxidizer (LOX) flow distribution through the LOX posts have a great effect on the combustion phenomena inside the main combustion chamber. In order to design a CFD model to be an effective engineering analysis tool with good computational turn- around time (especially for 3-D flow problems) and still maintain good accuracy in describing the flow features, the concept of porosity was employed to describe the effects of blockage and drag force due to the presence of the LOX posts in the turbulent flow field around the main injector assembly of the SSME. 2-D numerical studies were conducted to identify the drag coefficients of the flows both through tube banks and around the shielded posts over a wide range of Reynolds numbers. Empirical, analytical expressions of the drag coefficient as a function of local flow Reynolds number were then deduced. The porosity model was applied to the turbulent flow around the main injector assembly of the SSME, and analyses were performed. The 3-D CFD analysis was divided into three parts, LOX dome, hot gas injector assembly, and hydrogen cavity. The numerical results indicate that the mixture ratio at the downstream of injector face was close to stoichiometric around baffle elements.

Cheng, Gary C.↗
NASA NTRS1951

Arrangement of Bodies of Revolution in Supersonic Flow to Reduce Wave Drag

The wave drag of a combination of slender bodies of revolution at zero angle of attack is studied with a view to determining the arrangements for which the total drag is a minimum.Linearized theory is used to calculate the pressure distribution in the field surrounding the bodies. The interference drag coefficient is computed for different arrangements. The special cases of two bodies and of a three-body combination with bilateral symmetry are considered. The bodies treated are of the form determined by Sears and Haack as having minimum wave drag for prescribed volume and length. They also have equal fineness ratios. Numerical calculations of the drag coefficient of interference are carried out and curves are drawn which show the relative positions at which minimum drag occurs. A three-body configuration is found for which the total wave drag is about 35 percent less than the sum of the individual wave drags of the three bodies.

Friedman, Morris D↗
NASA NTRS1948

Effect of windshield shape of a pilot's canopy on the drag of an NACA RM-2 drag research model in flight at transonic speeds

Results of flight tests of an NACA RM-2 drag research model equipped with a pilot's canopy having a vee windshield are presented for a Mach number range from 0.75 to 1.43. Comparison is made with test results of a similar canopy having a flat windshield. The vee-windshield canopy produced lower drag-coefficient values than the flat-windshield canopy for Mach numbers from 0.85 to about 1.2. From M - 1.2 to 1.4 both canopies produced the same drag coefficient.

CANOPIES↗
NASA NTRS1967

Flight Test of 31.2 Diameter Modified Ringsail Parachute Deployed at Mach 1.39, Dynamic Pressure 11 Pounds per Square Foot

A 31.2-foot (9.51 meter) nominal diameter (reference area 764 ft(exp 2) (71.0 m(exp 2)) ringsail parachute modified to provide 15-percent geometric porosity was flight tested while attached to a 201-pound mass (91.2 kilogram) instrumented payload as part of the rocket launch portion of the NASA Planetary Entry Parachute Program (PEPP). The parachute deployment was initiated by the firing of a mortar at a Mach number of 1.39 and a dynamic pressure of 11.0 lb/ft(exp 2) (527 newtons/m(exp 2)) at an altitude of 122,500 feet (37.3 kilometers). The parachute deployed to suspension-line stretch (snatch force) in 0.35 second, and 0.12 second later the drag force increase associated with parachute inflation began. The parachute inflated in 0.24 second to the full-open condition for a total elapsed opening time of 0.71 second. The maximum opening load of 3970 pounds (17,700 newtons) came at the time the parachute was just fully opened. During the deceleration period, the parachute exhibited an average drag coefficient of 0.52 and oscillations of the parachute canopy were less than 5 degrees. During the steady-state terminal descent portion of the test period, the average effective drag coefficient (based on vertical descent velocity) was 0.52.

TESTS↗
NASA NTRS2016

Surface Stress in Tropical Cyclone Observed by Scatterometer

Ocean surface wind (U) is air in motion and stress (τ) is the turbulent transport of momentum between the ocean and the atmosphere. While the strong wind of a tropical cyclone (TC) causes destruction at landfall, it is the surface stress that drags down the TC. There was almost no stress measurement except in dedicated field campaigns and the stress we used was almost entirely derived from wind through a drag coefficient (C(sub D)), as defined by C(sub D) = τ / ( ρ U(exp 2)). In TC, there is difficulty in measuring strong wind and large uncertainty in the drag coefficient.

Liu, W. Timothy↗
NASA NTRS1962

DRAG CHARACTERISTICS AND DYNAMIC STABILITY IN DESCENT OF A ROTARY PARACHUTE TESTED IN A VERTICAL TUNNEL

An investigation was conducted in the Langley 20-foot free-spinning tunnel at low speeds to determine the drag characteristics and dynamic stability in descent of a 64-inch-diameter rotary parachute. The drag coefficient of the basic rotary parachute was approximately 1.7 based on total cloth area_ cutting cloth away from the tips of the parachute permitted it to rotate faster and gave a somewhat higher drag coefficient of about 2.1. In general, the rotary parachute was extremely stable in descent, with oscillations of probably less than 1°. Small changes in rigging had no appreciable effect on the performance of the parachute. Accompanying zip folder contains video of the experimental process completed for this document and the related records for the Experimental aerodynamic performance characteristics of a rotor entry vehicle configuration. 1-3.

WIND TUNNEL↗
NASA NTRS1987

Measurement of the effect of manufacturing deviations on natural laminar flow for a single engine general aviation airplane

Renewed interest in natural laminar flow (NLF) had rekindled designer concern that manufacuring deviations may destroy the effectiveness of NLF for an operational aircraft. Experiments are summarized that attemtped to measure total drag changes associated with three different wing surface conditions on an aircraft typical of current general aviation high performance singles. The speed power technique was first used in an attempt to quantify the changes in total drag. Predicted and measured boundary layer transition locations for three different wing surface conditions were also compared, using two different forms of flow visualization. The three flight test phases included: assessment of an unpainted airframe, flight tests of the same aircraft after painstakingly filling and sanding the wings to design contours, and similar measurement after this aricraft was painted. In each flight phase, transition locations were monitored using with sublimating chemicals or pigmented oil. Two-dimensional drag coefficients were estimated using the Eppler-Somers code and measured with a wake rake in a method very similar to Jones' pitot traverse method. The net change in two-dimensional drag coefficient was approximately 20 counts between the unpainted aircraft and the hand-smoothed aircraft for typical cruise flight conditions.

Source record↗
NASA NTRS1961

Aerodynamic Characteristics of Parachutes at Mach Numbers from 1.6 to 3

A wind-tunnel investigation has been conducted to determine the parameters affecting the aerodynamic performance of drogue parachutes in the Mach number range from 1.6 to 3. Flow studies of both rigid and flexible-parachute models were made by means of high-speed schlieren motion pictures and drag coefficients of the flexible-parachute models were measured at simulated altitudes from about 50,000 to 120,000 feet. Porosity and Mach number were found to be the most important factors influencing the drag and stability of flexible porous parachutes. Such parachutes have a limited range of stable'operation at supersonic speeds, except for those with very high porosities, but the drag coefficient decreases rapidly with increasing porosity.

Maynard, Julian D.↗
NASA NTRS1967

Flight Test of a 30-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at Mach 1.56 and Dynamic Pressure of 11.4 Pounds per Square Foot

A 30-foot (9.1 meter) nominal-diameter disk-gap-band parachute (reference area 707 sq ft (65.7 m(exp 2)) was flight tested with a 200-pound (90.7 kg) instrumented payload as part of the NASA Planetary Entry Parachute Program. A deployment mortar ejected the test parachute when the payload was at a Mach number of 1.56 and a dynamic pressure of 11.4 lb/sq ft (546 newtons per m 2 ) at an altitude of 127,500 feet (38.86 km). The parachute reached suspension line stretch in 0.37 second resulting in a snatch force loading of 1270 pounds (5650 N). Canopy inflation began 0.10 second after line stretch. A delay in the opening process occurred and was apparently due to a momentary interference of the glass-fiber shroud used in packing the parachute bag in the mortar. Continuous canopy inflation began 0.73 second after initiation of deployment and 0.21 second later full inflation was attained for a total elapsed time from mortar fire of 0.94 second. The maximum opening load of 3915 pounds (17,400 newtons) occurred at the time the canopy was first fully opened. The parachute exhibited an average drag coefficient of 0.52 during the deceleration period and pitch-yaw oscillations of the canopy were less than 5 degrees. During the steady-state descent portion of the test period, the average effective drag coefficient was about 0.47 (based on vertical descent velocity and total system weight).

EXPERIMENTS↗
NASA NTRS1961

Longitudinal Force and Moment Data at Mach Numbers from 0.60 to 1.40 for a Family of Elliptic Cones with Various Semiapex Angles

An investigation has been made to determine the aerodynamic characteristics of four elliptic cones having plan-form semiapex angles ranging from about 9 to 31 deg., and also for one of these cones modified on the upper surface to reduce the base area by about one half. The tests were made for angles of attack from about -2 to +21 deg., at Mach numbers from 0.60 to 1.40, and for a constant Reynolds number of 1.4 million, based on the length of the models. For each model, lift, pitching-moment, and drag coefficients, and lift-drag ratios are presented for the forebody, and axial-force coefficients are presented for the base. Calculated lift and pitching- moment curves for the elliptic cones, and lift-curve slopes for each model at supersonic Mach numbers are shown for comparison with the corresponding experimental values. Lift-drag ratios are also given for the forebody and base combined. These data are presented without discussion.

Stivers, Louis S., Jr.↗
NASA NTRS1952

An Experimental Investigation of Transonic Flow Past Two-Dimensional Wedge and Circular-Arc Sections Using A Mach-Zehnder Interferometer

Report presents the results of interferometer measurements of the flow field near two-dimensional wedge and circular-arc sections of zero angle of attack at high-subsonic and low-supersonic velocities. Both subsonic flow with local supersonic zone and supersonic flow with detached shock wave have been investigated. Pressure distributions and drag coefficients as a function of Mach number have been obtained. The wedge data are compared with the theoretical work on flow past wedge sections of Guderley and Yoshihara, Vincenti and Wagner, and Cole. Pressure distributions and drag coefficients for the wedge and circular-arc sections are presented throughout the entire transonic range of velocities.

Bryson, Arthur Earl, Jr↗
NASA NTRS1951

Effect of Tail Surfaces on the Base Drag of a Body of Revolution at Mach Numbers of 1.5 and 2.0

Wind-tunnel tests were performed at Mach numbers of 1.5 and 2.0 to investigate the influence of tail surfaces on the base drag of a body of revolution without boattailing and having a turbulent boundary layer. The tail surfaces were of rectangular plan form of aspect ratio 2.33 and has symmetrical, circular-arc airfoil section. The results of the investigation showed that the addition of these tail surfaces with the trailing edges at or near the body base incurred a large increase in the base-drag coefficient. For a cruciform tail having a 10-percent-thick airfoil section, this increase was about 70 percent at a Mach number of 1.5 and 35 percent at a Mach number of 2.0. As the trailing edge of the tail was moved forward or rearward of the base by about one tail-chord length, the base-drag increment was reduced to nearly zero. The increments in base-drag coefficient due to the presence of 10-percent-thick tail surfaces were generally twice those for 5-percent-thick surfaces. The base-drag increments due to the presence of a cruciform tail were less than twice those for a plane tail. An estimate of the change in base pressure due to the tail surfaces was made, based on a simple superposition of the airfoil-pressure field onto the base-pressure field behind the body. A comparison of the results with the experimental values indicated that in most cases the trend in the variation of the base-drag increment with changes in tail position could be predicted by this approximate method but that the quantitative agreement at most tail locations was poor.

BOUNDARY LAYER - BODIES OF REVOLUTION↗