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NASA NTRS1968

Flight Test of a 40-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 1.91 and a Dynamic Pressure of 11.6 Pounds per Square Foot

A 40-foot (12.2 meter) nominal-diameter disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator Program (SPED-I). The test parachute was ejected by a deployment mortar from an instrumented payload at an altitude of 140,000 feet (42.5 kilometers). The payload was at a Mach number of 1.91 and the dynamic pressure was 11.6 pounds per square foot (555 newtons per square meter) at the time the parachute deployment mortar was fired. The parachute reached suspension line stretch in 0.43 second with a resultant snatch force loading of 1990 pounds (8850 newtons). The maximum parachute opening load of 6500 pounds (28,910 newtons) came 0.61 second later at a total elapsed time from mortar firing of 1.04 seconds. The first full inflation occurred at 1.12 seconds and stable inflation was achieved at approximately 1.60 seconds. The parachute had an average axial-force coefficient of 0.53 during the deceleration period. During the steady-state descent portion of the flight test, the average effective drag coefficient was also 0.53 and pitch-yaw oscillations of the canopy averaged less than 10 degrees in the altitude region above 100,000 feet (30.5 meters).

EXPERIMENTS↗

NASA NTRS1957

Aerodynamic Characteristics of a 0.04956-Scale Model of the Convair TF-102A Airplane at Transonic Speeds, Coord. No. AF-120

The basic aerodynamic characteristics of a 0.04956-scale model of the Convair TF-102A airplane with controls undeflected have been determined at Mach numbers from 0.60 to 1.135 for angles of attack up to approximately 22 deg in the Langley 8-foot transonic tunnel. In addition, comparisons have been made with data obtained from a previous investigation of a 0.04956-scale model of the Convair F-102A airplane. The results indicated the TF-102A airplane was longitudinally stable for all conditions tested. An increase in lift-curve slope from 0.045 to 0.059 and an 11-percent rearward shift in aerodynamic-center location occurred with increases in Mach number from 0.60 to approximately 1.05. The zero-lift drag coefficient for the TF-102A airplane increased 145 percent between the Mach numbers of 0.85 and 1.075; the maximum lift-drag ratio decreased from 9.5 at a Mach number of 0.60 to 5.0 at Mach numbers above 1.025. There was little difference in the lift and pitching-moment characteristics and drag due to life between the TF-102A and F-102A configurations. However, as compared with the F-102A airplane, the zero-lift drag-rise Mach number for the TF-102A was reduced by at least 0.06, the zero-lift peak wave drag was increased 50 percent, and the maximum lift-drag ratio was reduced as much as 20 percent.

Osborne, Robert S.↗

NASA NTRS1992

A turbulence model for iced airfoils and its validation

A turbulence model based on the extension of the algebraic eddy viscosity formulation of Cebeci and Smith developed for two dimensional flows over smooth and rough surfaces is described for iced airfoils and validated for computed ice shapes obtained for a range of total temperatures varying from 28 to -15 F. The validation is made with an interactive boundary layer method which uses a panel method to compute the inviscid flow and an inverse finite difference boundary layer method to compute the viscous flow. The interaction between inviscid and viscous flows is established by the use of the Hilbert integral. The calculated drag coefficients compare well with recent experimental data taken at the NASA-Lewis Icing Research Tunnel (IRT) and show that, in general, the drag increase due to ice accretion can be predicted well and efficiently.

Shin, Jaiwon↗

NASA NTRS1992

A turbulence model for iced airfoils and its validation

A turbulence model based on the extension of the algebraic eddy viscosity formulation of Cebeci and Smith developed for two-dimensional flows over smooth and rough surfaces is described for iced airfoils and validated for computed ice shapes obtained for a range of total temperatures varying from 28 to -15 F. The validation is made with an interactive boundary layer method which uses a panel method to compute the inviscid flow and an inverse finite difference boundary layer method to compute the viscous flow. The interaction between inviscid and viscous flows is established by the use of the Hilbert integral. The calculated drag coefficients compare well with recent experimental data taken at the NASA-Lewis Icing Research Tunnel (IRT) and show that, in general, the drag increase due to ice accretion can be predicted well and efficiently.

Shin, Jaiwon↗

NASA NTRS1993

Simulated rarefied entry of the Galileo probe into the atmosphere of Jupiter

Flow properties and aerodynamics are computed with a direct simulation Monte Carlo (DSMC) method for rarefied entry of the Galileo Probe into the atmosphere of Jupiter. Accurate predictions of vehicle drag coefficients are needed in order to assess atmospheric properties from the onboard Atmosphere Structure Experiment where highly-sensitive accelerometers will measure the drag force to within 10-6 barr during the initial entry phase at high altitudes. The corresponding flow rarefraction extends from the free molecule limit to the near continuum transition regime (Re less than 1000). Simulation results indicate that C(sub D) varies from 2.1 at the free molecule limit down to 1.6 at Re(infinity) = 1,000. Temperatures, densities, and internal energies throughout the flow field were also computed at each altitude ranging from 735 km to 353 km above the 1 barr level in the Jovian atmosphere. Surface heating and temperatures of the probe were computed directly in the DSMC code by assuming radiative equilibrium. Material response was re-asssessed accurately during entry by accounting for conductivity, heat capacity, and pyrolysis which led to surface material mass efflux several times that of the freestream mass influx. The simulation also accounted for the quantum nature of the rotational energy mode of the dominant atmospheric species H2 through partial internal excitation in the freestream gas.

Hass, Brian L.↗

NASA NTRS2003

Numerical Simulations of the Steady and Unsteady Aerodynamic Characteristics of a Circulation Control Wing Airfoil

The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.

Liu, Yi↗

NASA NTRS1940

Preliminary Tests in the NACA Tank to Investigate the Fundamental Characteristics of Hydrofoils

This preliminary investigation was made to study the hydrodynamic properties and general behavior of simple hydrofoils. Six 5- by 30-inch plain, rectangular hydrofoils were tested in the NACA tank at various speeds, angles of attack and depths below the water surface. Two of the hydrofoils had sections representing the sections of commonly used airfoils, one had a section similar to one developed Guidoni for use with hydrofoil-equipped seaplane floats, and three had sections designed to have constant chordwise pressure distributions at given values of the lift coefficient for the purpose of delaying the speed at which cavitation begins. The experimental results are presented as curves of the lift and drag coefficients plotted against speed for the various angles of attack and depths for which the hydrofoils were tested. A number of derived curves are included for the purpose of better comparing the characteristics of the hydrofoils and to show the effects of depth. Several representative photographs show the development of cavitation on the the upper surface of the hydrofoils. The results indicate that properly designed hydrofoil sections will have excellent characteristics and that the speed at which cavitation occurs may be delayed to an appreciable extent by the use of suitable sections.

Ward, Kenneth E.↗

NASA NTRS1954

Flight Results from a 1/10-Scale Rocket Model of the Lockheed XF-104 Airplane at Transonic Mach Numbers

A 1/10-scale rocket model of the Lockheed XF-104 with faired inlets has been flown over a Mach number range from 0.80 to 1.45 to determine low-lift drag and a limited amount of stability data. The center-of-gravity locations were 4.0 and 1.5 percent of the mean aerodynamic chord before and after sustainer firing, respectively. Oscillations induced by pulse rockets were used to determine stability data. The external transonic drag coefficient increased from a value of 0.0160 at Mach number 0.80 to a maximum of 0.0432 near Mach number 1-13, with a drag rise Mach number of about 0.93. At Mach numbers where it could be determined, the model exhibited stable dynamic and static stability characteristics at low lift.

Kehlet, Alan B.↗

NASA NTRS1971

Drag of a Supercritical Body of Revolution in Free Flight at Transonic Speeds and Comparison with Wind Tunnel Data

The forebody drag of a supercritical body of revolution was measured in free flight over a Mach number range of 0.85 to 1.05 and a Reynolds number range of 11.5 x 10 to the 6th power to 19.4 x 10 to the 6th power and was compared with wind-tunnel data. The forebody drag coefficient for a Mach number less than 0.96 was 0.111 compared with the wind-tunnel value of 0.103. A gradual increase in the drag occurred in the Langley 8-foot transonic pressure tunnel at a lower Mach number than in the Langley 16-foot transonic tunnel or in the free-flight test. The sharp drag rise occurred near Mach 0.98 in free flight whereas the rise occurred near Mach 0.99 in the Langley 16-foot transonic tunnel. The sharp rise was not as pronounced in the Langley 8-foot transonic pressure tunnel and was probably affected by tunnel-wall-interference effects. The increase occurred more slowly and at a higher Mach number. These results indicate that the drag measurements made in the wind tunnels near Mach 1 were significantly affected by the relative size of the model and the wind tunnel.

Usry, J. W.↗

NASA NTRS1928

Drag of Exposed Fittings and Surface Irregularities on Airplane Fuselages

Measurements of drag were made on fittings taken from a typical fuselage to determine whether the difference between the observed full size fuselage drag and model fuselage drag could be attributed to the effects of fittings and surface irregularities found on the full size fuselage and not on the model. There are wide variations in the drag coefficients for the different fittings. In general those which protrude little from the surface or are well streamlined show very low and almost negligible drag. The measurements show, however, that a large part of the difference between model and full scale test results may be attributed to these fittings.

Wood, Donald H↗

NASA NTRS1950

Free-Flight Investigation of the Static and Dynamic Longitudinal Stability Characteristics of 1/3.7-Scale Rocket-Powered Models of the Bell MX-776A

An investigation of the static and dynamic longitudinal stability characteristics of 1/3.7 scale rocket-powered model of the Bell MX-776A has been made for a Mach number range from 0.8 to 1.6. Two models were tested with all control surfaces at 0 degree deflection and centers of gravity located 1/4 and 1/2 body diameters, respectively, ahead of the equivalent design location. Both models were stable about the trim conditions but did not trim at 0 degree angle of attack because of slight constructional asymmetries. The results indicated that the variation of lift and pitching moment was not linear with angle of attack. Both lift-curve slope and pitching-moment-curve slope were of the smallest magnitude near 0 degree angle of attack. In general, an increase in angle of attack was accompanied by a rearward movement of the aerodynamic center as the rear wing moved out of the downwash from the forward surfaces. This characteristic was more pronounced in the transonic region. The dynamic stability in the form of total damping factor varied with normal-force coefficient but was greatest for both models at a Mach number of approximately 1.25. The damping factor was greater at the lower trim normal-force coefficients except at a Mach number of 1.0. At that speed the damping factor was of about the same magnitude for both models. The drag coefficient increased with trim normal-force coefficient and was largest in the transonic region.

Michal, David H.↗

NASA NTRS1927

Pressure Distribution Over Airfoils at High Speeds

This report deals with the pressure distribution over airfoils at high speeds, and describes an extension of an investigation of the aerodynamic characteristics of certain airfoils which was presented in NACA Technical Report no. 207. The results presented in report no. 207 have been confirmed and extended to higher speeds through a more extensive and systematic series of tests. Observations were also made of the air flow near the surface of the airfoils, and the large changes in lift coefficients were shown to be associated with a sudden breaking away of the flow from the upper surface. The tests were made on models of 1-inch chord and comparison with the earlier measurements on models of 3-inch chord shows that the sudden change in the lift coefficient is due to compressibility and not to a change in the Reynolds number. The Reynolds number still has a large effect, however, on the drag coefficient. The pressure distribution observations furnish the propeller designer with data on the load distribution at high speeds, and also give a better picture of the air-flow changes.

Briggs, L J↗

NASA NTRS1950

Comparative Drag Measurements at Transonic Speeds of Rectangular Sweptback NACA 65-009 Airfoils Mounted on a Freely Falling Body

Directly comparable drag measurements have been made of an airfoil with a conventional rectangular plan form and an airfoil with a sweptback plan form mounted on freely falling bodies. Both airfoils had NACA 65-009 sections and were identical in span, frontal area, and chord perpendicular to the leading edge. The sweptback plan form incorporated a sweepback angle of 45 degrees. The data obtained have been used to establish the relation between the airfoil drag coefficients and the free-stream Mach number over a range of Mach numbers from 0.90 to 1.27. The results of the measurements indicate that the drag of the sweptback plan form is less than 0.3 that of the rectangular plan form at a Mach number of 1.00 and is less than 0.4 that at a Mach number of 1.20.

Mathews, Charles W↗

NASA NTRS1947

Development of Outboard Nacelle for the XB-36 Airplane

An investigation of two 1/14 scale model configurations of an outboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels over a range of airplane lift coefficients (C (sub L) = 0.409 to C(sub L) = 0.943) for three representative flow conditions. The purpose of the investigation was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. The present investigation has led to the development of a nacelle which had external drag coefficients of similar order of magnitude to those obtained previously from tests of an inboard nacelle configuration at the corresponding operating lift coefficients and from approximately one-third to one-half of those of conventional tractor designs having the same ratio of wing thickness to nacelle diameter.

Nuber, Robert J.↗

NASA NTRSDate not supplied

Sensitivity Study of Dust-Induced Surface Erosion during Martian Planetary Entry

Large, global dust storms that periodically form in the Martian atmosphere pose a threat to NASA missions due to a lack of understanding of how they will affect the heat loading and mass loss of the entry vehicle’s thermal protection system (TPS). Previous studies have laid out viable methodologies for estimating the amount of surface recession from a significant dust storm, but rely on underlying modeling assumptions that are difficult to validate. The current work utilizes the DUST library, which is a Lagrangian particle solver under development at NASA Ames Research Center. DUST includes variable-fidelity fluid-particle coupling and has been tested for a wide range of problem types. The DUST library is used in conjunction with the US3D Naiver-Stokes solver. Using these tools, a sensitivity study is performed to determine the influence of the underlying dust modeling (drag coefficient, Nusselt number, and size distribution) on the surface erosion through a hypersonic Martian entry trajectory. The surface erosion was found to be insensitive to the drag and heating models, but was affected significantly by the size distributions parameters. The total recession was comparable to the thermochemical ablation predicted by material response codes, but the lack of a particle erosion model for PICA limits the predictive capability of the current work.

Thermal Protection Systems↗

NASA NTRS1978

Effect of camber on the trimmed lift capability of a close-coupled canard-wing configuration

A close-coupled canard-wing configuration was tested in the Langely high-speed 7 by 10 foot tunnel at a Mach number of 0.30 to determine the effect of changing wing camber on the trimmed lift capability. Trimmed lift coefficients of near 2.0 were attained; however, the data indicated that the highest buffet-free trimmed lift coefficient attainable was approximately 1.30. The buffet used in this investigation were qualitative in nature and gave no indication of buffet intensity. Thus, the trimmed lift coefficient of near 2.0 might be attainable if the buffet intensity was not too high. The data showed that there was approximately a 10 percent variation in drag coefficient, for different model configurations, at a given trimmed lift coefficient. Large increases in wing lift had only small effects on canard lift.

Gloss, B. B.↗

NASA NTRS1951

Aerodynamic Characteristics of a Flying-Boat Hull Having a Length-Beam Ratio of 15, TED No. NACA 2206

An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the aerodynamic characteristics of a flying-boat hull of a length-beam ratio of 15 in the presence of a wing. The investigation was an extension of previous tests made on hulls of length-beam ratios of 6, 9, and 12; these hulls were designed to have approximately the same hydrodynamic performance with respect to spray and resistance characteristics. Comparison with the previous investigation at lower length-beam ratios indicated a reduction in minimum drag coefficients of 0.0006 (10 peroent)with fixed transition when the length-beam ratio was extended from 12 to 15. As with the hulls of lower length-beam ratio, the drag reduction with a length-beam ratio of 15 occurred throughout the range of angle of attack tested and the angle of attack for minimum drag was in the range from 2deg to 3deg. Increasing the length-beam ratio from 12 to 15 reduced the hull longitudinal instability by an mount corresponding to an aerodynamic-center shift of about 1/2 percent of the mean aerodynamic chord of the hypothetical flying boat. At an angle of attack of 2deg, the value of the variation of yawing-moment coefficient with angle of yaw for a length-beam ratio of 15 was 0.00144, which was 0.00007 larger than the value for a length-beam ratio of 12.

Riebe, John M.↗

NASA NTRS1954

Theoretical and experimental investigation of additive drag

The significance of additive drag is discussed and equations for determining its approximate value are derived for annular and open-nose inlets. Charts are presented giving values of additive drag coefficient over a range of free-stream Mach numbers for open and for annular-nose inlets with conical flow at the inlet. The effects on additive drag of variable inlet-total-pressure recovery and static pressures on the centerbody are investigated and an analytical method of predicting the variation of pressure on the centerbody with mass-flow ratio is given. Experimental additive-drag values are presented for a series of 20 degree and 25 degree cone half-angle inlets and one open-nose inlet operating at free-stream Mach numbers of 1.8 and 1.6. A comparison with the theoretical values of additive drag shows excellent agreement for the open-nose inlet and moderately good agreement for the annular inlets. (author)

Sibulkin, Merwin↗