Results for “AFTS” — Page 2

NASA NTRS1976

SRM attrition rate study of the aft motor case segments due to water impact cavity collapse loading

The attrition assessment of the aft segments of Solid Rocket Motor due to water impact requires the establishment of a correlation between loading occurrences and structural capability. Each discrete load case, as identified by the water impact velocities and angle, varies longitudinally and radially in magnitude and distribution of the external pressure. The distributions are further required to be shifted forward or aft one-fourth the vehicle diameter to assure minimization of the effect of test instrumentation location for the load determinations. The asymmetrical load distributions result in large geometric nonlinearities in structural response. The critical structural response is progressive buckling of the case. Discrete stiffeners have been added to these aft segments to aid in gaining maximum structural capability for minimum weight addition for resisting these loads. This report presents the development of the attrition assessment of the aft segments and includes the rationale for eliminating all assessable conservatisms from this assessment.

Crockett, C. D.↗

NASA NTRS1980

A Study of Panel Loads and Centers of Pressure of Three Different Cruciform Aft-Tail Control Surfaces of a Wingless Missile from Mach 1.60 to 3.70

An investigation was made of the forces and moments on the cruciform aft-tail control surfaces of a wingless missile model to determine the variation of panel load and center of pressure with angle of attack, tail deflection, model roll angle, and Mach number. Also, a limited force-moment and surface-pressure investigation was made on a noncircular aft end. These investigations were made in a unitary plan wind tunnel at Mach numbers of 1.60, 2.36, and 3.70 and at a Reynolds number per meter of 6,600,000. The cruciform aft-tail results indicate very little variation in the center of pressure for the highly loaded windward tail. The noncircular aft-end test results indicate no significant effect of the fin unporting on the fin loads.

Lamb, M.↗

NASA NTRS1986

Static internal performance of a thrust vectoring and reversing two-dimensional convergent-divergent nozzle with an aft flap

The static internal performance of a multifunction nozzle having some of the geometric characteristics of both two-dimensional convergent-divergent and single expansion ramp nozzles has been investigated in the static-test facility of the Langley 16-Foot Transonic Tunnel. The internal expansion portion of the nozzle consisted of two symmetrical flat surfaces of equal length, and the external expansion portion of the nozzle consisted of a single aft flap. The aft flap could be varied in angle independently of the upper internal expansion surface to which it was attached. The effects of internal expansion ratio, nozzle thrust-vector angle (-30 deg. to 30 deg., aft flap shape, aft flap angle, and sidewall containment were determined for dry and afterburning power settings. In addition, a partial afterburning power setting nozzle, a fully deployed thrust reverser, and four vertical takeoff or landing nozzle, configurations were investigated. Nozzle pressure ratio was varied up to 10 for the dry power nozzles and 7 for the afterburning power nozzles.

Re, R. J.↗

NASA NTRS1987

Intergration effects of D-shaped, underwing, aft-mounted, separate-flow, flow-through nacelles on a high-wing transport

An experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at freestream Mach numbers from 0.70 to 0.82 and angles of attack from -3.0 to 4.0 deg to determine the integration effects of D-shaped, underwing, aft-mounted, separate-flow, flow-through nacelles on a high-wing transonic transport configuration. The results showed that the aft-mounted nacelle/pylon produced an increase in lift over that of the wing-body configuration by pressurizing much of the wing lower surface in front of the pylon. For the D-shaped nacelle, a substantial region of supersonic flow over the wing, aft of the lip of the nacelle, cancelled the reduction in drag caused by the increase in pressures ahead of the lip, to increase interference and form drag compared with a similar circular-shaped nacelle. The installed drag of the D=shaped nacelle was essentially the same as that of an aft-mounted circular nacelle from a previous investigation.

Lamb, Milton↗

NASA NTRS1988

Noise of a model counterrotation propeller with reduced aft rotor diameter at simulated takeoff/approach conditions (F7/A3)

A model high-speed advanced counterrotation propeller, F7/A3, was tested in the NASA Lewis Research Center 9 by 15 foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach number. Acoustic measurements were taken with an axially translating microphone probe, and with a polar microphone probe which was fixed to the propeller nacelle and could take both sideline and circumferential acoustic surveys. Aerodynamic measurements were also made to establish propeller operating conditions. The propeller was run at two setting angles (front angle/rear angle) of 36.4/43.5 and 41.1/46.4 degrees, forward rotor tip speeds from 165 to 259 m/sec, rotor spacings from 8.48 to 14.99 cm based on pitch change axis separation, and angles of attack to 16 degrees. The aft rotor diameter was 85 percent of the forward rotor diameter to reduce tip vortex-aft rotor interaction as a major interaction noise source. Results are compared with equal diameter F7/A7 data which was previously obtained under similar operating conditions. The aft rotor-alone tone was 7 dB lower for the reduced diameter aft rotor, due to reduced tip speed at constant rpm. Interaction tone levels for the F7/A3 propeller were higher at minimum row spacing and lower at maximum spacing.

Woodward, Richard P.↗

NASA NTRS1988

Noise of a model counterrotation propeller with reduced aft rotor diameter at simulated takeoff/approach conditions (F7/A3)

A model high-speed advanced counterrotation propeller, F7/A3, was tested in the NASA Lewis Research Center 9 by 15 foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach number. Acoustic measurements were taken with an axially translating microphone probe, and with a polar microphone probe which was fixed to the propeller nacelle and could take both sideline and circumferential acoustic surveys. Aerodynamic measurements were also made to establish propeller operating conditions. The propeller was run at two setting angles (front angle/rear angle) of 36.4/43.5 and 41.1/46.4 degrees, forward rotor tip speeds from 165 to 259 m/sec, rotor spacings from 8.48 to 14.99 cm based on pitch change axis separation, and angles of attack to 16 degrees. The aft rotor diameter was 85 percent of the forward rotor diameter to reduce tip vortex-aft rotor interaction as a major interaction noise source. Results are compared with equal diameter F7/A7 data which was previously obtained under similar operating conditions. The aft rotor-alone tone was 7 dB lower for the reduced diameter aft rotor, due to reduced tip speed at constant rpm. Interaction tone levels for the F7/A3 propeller were higher at minimum row spacing and lower at maximum spacing.

Woodward, Richard P.↗

NASA NTRS1989

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 1: Aft Skirt analysis

Using the ANSYS finite element program, a global model of the aft skirt and a detailed nonlinear model of the failure region was made. The analysis confirmed the area of failure in both STA-2B and STA-3 tests as the forging heat affected zone (HAZ) at the aft ring centerline. The highest hoop strain in the HAZ occurs in this area. However, the analysis does not predict failure as defined by ultimate elongation of the material equal to 3.5 percent total strain. The analysis correlates well with the strain gage data from both the Wyle influence test of the original design aft sjirt and the STA-3 test of the redesigned aft skirt. it is suggested that the sensitivity of the failure area material strength and stress/strain state to material properties and therefore to small manufacturing or processing variables is the most likely cause of failure below the expected material ultimate properties.

Berry, David M.↗

NASA NTRS1989

Effect of reduced aft diameter and increased blade number on high-speed counterrotation propeller performance

Performance data of 0.17-scale model counterrotation pusher propeller configurations were taken in the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel at Mach numbers of 0.66, 0.71, 0.75, and 0.79. These tests investigated the aerodynamic performance of the unducted fan (UDF) demonstrator propeller engine developed in a joint program by General Electric and NASA. Data were recorded to show the effect on counterrotation propeller cruise efficiency of two takeoff noise-reduction concepts. These two concepts are reduced aft blade diameter and increased forward blade number. The four configurations tested were a baseline (F1/A1 8/8) configuration, a reduced aft diameter (F1/A3 8/8) configuration, an increase forward blade number (F1/A1 9/8) configuration, and a combination of the latter two (F1/A3 9/8) configurations. Data were collected with a complex counterrotation propeller test rig via rotating thrust and torque balances and pressure instrumentation. Data comparisons documented the power differences between the baseline and the reduced aft diameter concepts. Performance comparisons to the baseline configuration showed that reducing the aft blade diameter reduced the net efficiency, and adding a blade to the front rotor increased the net efficiency. The combination of the two concepts showed only slightly lower net efficiency than the baseline configuration. It was also found that the counterrotation demonstrator propeller model (F7/A7 8/8) configuration outperformed the baseline (F1/A1 8/8) configuration.

Rose, Gayle E.↗

NASA NTRS1989

Effect of reduced aft diameter and increased blade number of high-speed counterrotation propeller performance

Performance data of 0.17-scale model counterrotation pusher propeller configurations were taken in the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel at Mach numbers of 0.66, 0.71, 0.75, and 0.79. These tests investigated the aerodynamic performance of the unducted fan (UDF) demonstrator propeller engine developed in a joint program by General Electric and NASA. Data were recorded to show the effect on counterrotation propeller cruise efficiency of two takeoff noise-reduction concepts. These two concepts are reduced aft blade diameter and increased forward blade number. The four configurations tested were a baseline (F1/A1 8/8) configuration, a reduced aft diameter (F1/A3 8/8) configuration, an increase forward blade number (F1/A1 9/8) configuration, and a combination of the latter two (F1/A3 9/8) configurations. Data were collected with a complex counterrotation propeller test rig via rotating thrust and torque balances and pressure instrumentation. Data comparisons documented the power differences between the baseline and the reduced aft diameter concepts. Performance comparisons to the baseline configuration showed that reducing the aft blade diameter reduced the net efficiency, and adding a blade to the front rotor increased the net efficiency. The combination of the two concepts showed only slightly lower net efficiency than the baseline configuration. It was also found that the counterrotation demonstrator propeller model (F7/A7 8/8) configuration outperformed the baseline (F1/A1 8/8) configuration.

Gayle, E. Rose↗

NASA NTRS1997

Acoustic Barrier Facilitates Inlet Noise Measurements for Aft-Dominated Fans

Noise levels for modern high-bypass-ratio subsonic turbofans tend to be aft dominated. That is, the highest flyover noise levels radiate from the fan exit. Measuring fan inlet sound radiation without aft radiation contamination requires selective suppression of the aft noise. In NASA Lewis Research Center's 9- by 15-Foot Low-Speed Wind Tunnel, an acoustic barrier was used to effectively isolate the inlet noise field for a model of an advanced turbofan. This proof-of-concept test was performed on a model turbofan manufactured for NASA Lewis by the Allison Engine Company as part of the Advanced Subsonic Technology program. The 8-cm-thick acoustic barrier was constructed in sections that were joined upon installation. These sections, which were composed of a wood frame with typically 0.64-cm tempered fiberboard skins, extended from the tunnel's floor to its ceiling and had an axial length of 61 cm. On the fan side of the barrier just downstream of the leading edge, the upstream section had an acoustic treatment--a bulk absorber with a perforated metal skin. It had a nominal full height and an axial length of 46 cm. In addition, an elliptical leading edge was faired into the upstream barrier section. The barrier was mounted on tracks on the tunnel floor and ceiling at a sideline distance of 15 cm from the fan nacelle. Tests were made with the barrier leading edge at the fan inlet highlight plane and 15 cm further aft. The barrier extended downstream essentially to the end of the treated tunnel test section.

Source record↗

NASA NTRS2013

Enabling a Better Aft Heat Shield Solution for Future Mars Science Laboratory Class Vehicles

System studies are described that compare masses and estimated manufacturing costs of options for the as‐flown Mars Science Laboratory (MSL) aft body Thermal Light Weight Ablator (SLA) 561‐V and its thickness was not optimized using the standard TPS Sizer Tool widely used for heat shield design. Use of the TPS sizing tool suggests that optimization of the SLA thickness could reduce the aft heat shield mass by 40 percent. Analysis of the predicted aft‐shell aerothermodynamics suggests that the bulk of MSL class entry vehicle heat shields could incorporate Advanced Flexible Reusable Surface Insulation (AFRSI). AFRSI has a wellestablished record of relatively inexpensive manufacturing and flight certification based on its use on the lee side of the Space Shuttle. Runs with the TPS Sizer show that the AFRSI solution would be 60 percent lighter than the as‐flown SLA. The issue of Reaction Control System (RCS) heating on the aft shell could be addressed by locally impregnating the AFRSI with silicone to enhance its robustness to short bursts ofheating. Stagnation point arcjet testing has shown that silicone impregnated AFRSI performs well at heat rates of 115 W/cm2 and 0.1 atmospheres for a duration of 40 seconds, far beyond conditions that are expected for MSL class vehicles. The paper concludes with a discussion of manufacturing processes for AFRSI, impregnation approaches and relative cost comparisons to the SLA solution.

McGuire, Mary K.↗

NASA NTRS1975

Integration of aft-fuselage-mounted flow through engine nacelles on an advanced transport configuration at Mach numbers from 0.6 to 1.0

An approximately 1/30th-scale model of a CTOL aircraft designed for efficient cruise between Mach 0.9 and 0.98 was tested in the Lewis Research Center's 8- by 6-Foot Supersonic Wind Tunnel. Aft-fuselage axial force and pressure were measured for a series of locally area-ruled engine installations. Two flow-through nacelles with different exit diameters which simulated inlet mass flow ratios of 0.46 and 0.6 (based on maximum nacelle cross-sectional) area were investigated. Both an NACA-1 cowl and a blunter supercritical cowl were tested with these nacelles. Variations in the amount of local area ruling were investigated and an aft fuselage without nacelles was tested for reference. Results observed with flow-through nacelles indicate that efficient cruise may be obtained on this type of aircraft with aft-fuselage-mounted engine nacelles.

Blaha, B. J.↗

NASA NTRS1976

SRB attrition rate study of the aft skirt due to water impact cavity collapse loading

A methodology was presented so that realistic attrition prediction could aid in selecting an optimum design option for minimizing the effects of updated loads on the Space Shuttle Solid Rocket Booster (SRB) aft skirt. The updated loads resulted in water impact attrition rates greater than 10 percent for the aft skirt structure. Adding weight to reinforce the aft skirt was undesirable. The refined method treats the occurrences of the load distribution probabilistically, radially and longitudinally, with respect to the critical structural response.

Crockett, C. D.↗

NASA NTRS1979

Space Shuttle Orbiter AFT heat shield seal

The evolution of the orbiter aft heat shield seal (AHSS) design, which involved advancing mechanical seal technology in severe thermal environment is discussed. The baseline design, various improvements for engine access, and technical problem solution are presented. It is a structure and mechanism at the three main propulsion system (MPS) engine interfaces to the aft compartment structure. Access to each MPS engine requires disassembly and removal of the AHSS. Each AHSS accommodates the engine movement, is exposed to an extremely high temperature environment, and is part of the venting control of the aft compartment.

Walkover, L. J.↗

NASA NTRS1980

Far-field radiation of aft turbofan noise

Approximate expressions were developed for the noise radiation from the aft duct. The results of approximate aft radiation equation compare favorably to more exact Wiener-Hopf radiation results. Refraction as well as convective effects in the multiple flow streams is considered. The peak in the radiation pattern, which occurs nearly at engine sideline, is composed of modes with relatively large cut-off ratios. This implies that aft fan radiation will be inherently more difficult to suppress that the fan inlet noise. The theoretical multimodal radiation pattern is compared to experimental data for the first two harmonics of blade passage frequency for three full scale fans at two speeds. The agreement between theory and experiment is quite good.

Rice, E. J.↗

NASA NTRS1983

Interference effects of aft reaction-control yaw jets on the aerodynamic characteristics of a space shuttle orbiter model at supersonic speeds

A wind tunnel investigation of the interference effects of aft reaction control system yaw jet plumes on a 0.0125 scale Space Shuttle orbiter model was conducted at Mach numbers from 2.50 to 4.50. Test variables included model angle of attack, model angle of sideslip, jet to free stream mass flow ratio, and number and position of operating jets. The aft reaction control jet plume creates a blockage above and behind the wing on the side in which the jet exhausts and results in flow separation on the wing upper surface and fuselage side. Positive pitching moment and side force increments and negative yawing moment and rolling moment increments due to the flow separations are incurred for left side firing jets, primarily at angles of attack above 10 deg. The yawing moment interference increments are favorable and result in a small jet thrust amplification. As a result of this investigation, the aft reaction control system was certified for operation at supersonic Mach numbers prior to the first flight of the space transportation system (STS-1).

Covell, P. F.↗

NASA NTRS1984

Space Shuttle with rail system and aft thrust structure securing solid rocket boosters to external tank

The configuration and relationship of the external propellant tank and solid rocket boosters of space transportation systems such as the space shuttle are described. The space shuttle system with the improved propellant tank is shown. The external tank has a forward pressure vessel for liquid hydrogen and an aft pressure vessel for liquid oxygen. The solid rocket boosters are joined together by a thrust frame which extends across and behind the external tank. The thrust of the orbiter's main rocket engines are transmitted to the aft portion of the external tank and the thrust of the solid rocket boosters are transmitted to the aft end of the external tank.

Vonpragenau, G. L.↗

NASA NTRS1985

Supersonic aerodynamic characteristics of canard, tailless, and aft-tail configurations for 2 wing planforms

Aerodynamic characteristics of canard, tailless, and aft tail configurations were compared in tests on a general research model (generic fuselage without canopy, inlets, or vertical tails) at Mach 1.60 and 2.00 in the Langley Unitary Plan Wind Tunnel. Two uncambered wing planforms (trapezoidal with 44 deg leading edge sweep and delta with 60 deg leading edge sweep) were tested for each configuration. The relative merits of the configurations were also determined theoretically, to evaluate the capabilities of a linear theory code for such analyses. The canard and aft tail configurations have similar measured values for lift curve slope, maximum lift drag ratio, and zero lift drag. The stability decrease as Mach number increases is greatest for the tailless configuration and least for the canard configuration. Because of very limited accuracy in predicting the aerodynamic parameter increments between configurations, the linear theory code is not adequate for determining the relative merits of canard, tailless, and aft tail configurations.

Covell, P. F.↗