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  • 1
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    Low speed test of a high-bypass-ratio propulsion system with an asymmetric inlet designed for a tilt-nacelle V/STOL airplane (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A large scale model of a lift/cruise fan inlet designed for a tilt nacelle V/STOL airplane was tested with a high bypass ratio turbofan. Testing was conducted at low freestream velocities with inlet angles of attack ranging from 0 deg to 120 deg. The operating limits for the nacelle were found to be related to inlet boundary layer separation. Small separations originating in the inlet diffuser cause little or no performance degradation. However, at sufficiently severe freestream conditions the separation changes abruptly to a lip separation. This change is associated with a significant reduction in nacelle net thrust as well as a sharp increase in fan blade vibratory stresses. Consequently, the onset of lip separation is regarded as the nacelle operating limit. The test verified that the asymmetric inlet design will provide high performance and stable operation at the design forward speed and angle of attack conditions. At some of these, however, operation near the lower end of the design inlet airflow range is not feasible due to the occurrence of lip separation.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-152072 , D-180-22888-1
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Transonic and supersonic test of a Mach 2.65 mixed-compression axisymmetric intake (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The test results describe isolated intake performance between Mach 0.95 and the cruise Mach number of 2.65 at angles of incidence from +5 to -5 deg. Maximum total pressure recoveries of over 94 percent with 10 percent distortion were recorded at the compressor face in the Mach range from 2.65 to 2.4. Typical cruise operating recovery was 91 percent with 13 percent distortion, 7 percent bleed, 5 percent corrected flow stability margin, and 2.2 deg angle-of-incidence tolerance without need for control action. In the started range below Mach 2.4, recoveries were 2 percent to 4 percent lower than the recoveries above Mach 2.4, and the distortion increased to approximately 20 percent. At Mach 0.95 the maximum measured capture flow was 99.4 percent of the theoretical choked value. The recovery was 97.1 percent with less than 10 percent distortion.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-1977
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Aerodynamic analysis of VTOL inlets and definition of a short, blowing-lip inlet (1982)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results indicated that, without boundary layer control, either a very long inlet or an inlet with a very high contraction ratio lip will be required to meet the stringent design requirements. It is shown that active boundary layer control is an effective means of preventing separation and that a significant reduction in inlet size can be achieved by removing only a small amount of bleed in the throat region of the inlet. A short, blowing-lip model was designed and fabricated. This model features an adjustable, blowing slot located near the hilite on the windward side of the inlet.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165617 , NAS 1.26:165617 , D6-51418
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Experimental evaluation of a Mach 3.5 axisymmetric inlet (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Wind tunnel test results for a large scale inlet model designed for Mach 3.5 are presented and compared with analytical predictions. The inlet is an axisymmetric mixed-compression type with a lip diameter of 49.723 cm. The inlet design was developed using analytical procedures. Data are shown for freestream Mach numbers from 0.6 to 3.5. The test results indicate that boundary layer bleed requirements can be accurately predicted. Good agreement was obtained with analytical predictions of the flowfield structure and boundary layer development in the supersonic diffuser yielding high performance at the design Mach number. The highest engine face total pressure recovery at Mach 3.5 was 85.8%; this was obtained at 0.05 Mach tolerance with only 2.8% total pressure distortion and 13.4% bleed. In the started Mach number range from 1.6 to 3.5, the total pressure recovery in the throat, downstream of the terminal normal shock, ranged between 91% and 95%. Total pressure losses in the subsonic diffuser varied from 3% to 13%. The higher losses occuring between Mach 2.5 and 3.2 were believed to be caused by the rapid rate of increase in the area of the diffuser just downstream of the throat, possibly coupled with inadequate centerbody throat bleed. In the unstarted mode at transonic speeds, the maximum inlet flow was over 99% of the theoretical maximum capture mass-flow.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-2563 , D6-42494
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Reduction of acoustic disturbances in the test section of supersonic wind tunnels by laminarizing their nozzle and test section wall boundary layers by means of suction (1974)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The feasibility of quiet, suction laminarized, high Reynolds number (Re) supersonic wind tunnel nozzles was studied. According to nozzle wall boundary layer development and stability studies, relatively weak area suction can prevent amplified nozzle wall TS (Tollmien-Schlichting) boundary layer oscillations. Stronger suction is needed in and shortly upstream of the supersonic concave curvature nozzle area to avoid transition due to amplified TG (Taylor-Goertler) vortices. To control TG instability, moderately rapid and slow expansion nozzles require smaller total suction rates than rapid expansion nozzles, at the cost of larger nozzle length Re and increased TS disturbances. Test section mean flow irregularities can be minimized with suction through longitudinal or highly swept slots (swept behind local Mach cone) as well as finely perforated surfaces. Longitudinal slot suction is optimized when the suction-induced crossflow velocity increases linearly with surface distance from the slot attachment line toward the slot (through suitable slot geometry). Suction in supersonic blowdown tunnels may be operated by one or several individual vacuum spheres.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: NASA-CR-2436 , D6-41564
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Fan inlet for a V/STOL airplane (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: An inlet was designed for a tilt nacelle lift/cruise fan. A complete set of low speed data was obtained with a 38-cm diameter cold-flow model, and a partial set of data was obtained with a 140-cm-diameter inlet coupled to a high bypass turbofan. The results indicate that a fixed-geometry inlet is a viable concept for the proposed application. Analyses of the data demonstrate that at high angles of incidence and at low airflow rates inlet-flow separation is strongly sensitive to Reynolds number and, consequently, the large-scale inlet has a significantly greater range of separation-free operation.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AIAA PAPER 77-802 , Propulsion Conference; Jul 11, 1977 - Jul 13, 1977; Orlando, FL; US
    Format: text
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    NASA TECHNICAL REPORTS
  • 7
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    Low speed and angle of attack effects on sonic and near-sonic inlets (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Tests of the Quiet, Clean Short-Haul Experimental Engine (QCSEE) were conducted to determine the effects of forward velocity and angle of attack on sonic and near-sonic inlet aerodynamic performance penalties and acoustic suppression characteristics. The tests demonstrate that translating centerbody and radial vane sonic inlets, and QCSEE high throat Mach number inlets, can be designed to operate effectively at forward speed and moderate angle of attack with good performance and noise suppression capability. The test equipment and procedures used in conducting the evaluation are described. Results of the tests are presented in tabular form.
    Keywords: AERODYNAMICS
    Type: NASA-CR-134778 , D6-42392
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Design of a bleed system for a Mach 3.5 inlet (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An analytical bleed system design procedure is applied to a Mach 3.5, axisymmetric, translating centerbody inlet. The design features a traveling centerbody bleed system with a ducting arrangement separating low-, medium-, and high-pressure bleed. Three forward plenums and nine throat plenums are installed on the centerbody to meet the bleed requirements in the started Mach range between 1.6 and 3.5. The cowl contains four stationary plenums, three for forward and one for throat bleed. To achieve maximum bleed plenum pressure and thereby minimize bleed drag all bleed holes are inclined 20 deg to the surface except in the cowl throat region. Here the requirement of high stability margin with minimum total pressure recovery penalty resulted in 90 deg bleed holes. The bleed hole diameter varies from bleed plenum to bleed plenum to achieve the most efficient boundary layer control, while the bleed exits are sized to operate at the highest possible plenum pressure without unchoking the bleed holes. Bleed flow rates, bleed plenum pressures, and boundary layer development along the cowl and centerbody are predicted over the entire started Mach range.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-2187 , D6-60168
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Low speed tests of a fixed geometry inlet for a tilt nacelle V/STOL airplane (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Test data were obtained with a 1/4 scale cold flow model of the inlet at freestream velocities from 0 to 77 m/s (150 knots) and angles of attack from 45 deg to 120 deg. A large scale model was tested with a high bypass ratio turbofan in the NASA/ARC wind tunnel. A fixed geometry inlet is a viable concept for a tilt nacelle V/STOL application. Comparison of data obtained with the two models indicates that flow separation at high angles of attack and low airflow rates is strongly sensitive to Reynolds number and that the large scale model has a significantly improved range of separation-free operation.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-CR-151922 , D180-20276-1
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Inlet design technology development - Supersonic cruise research (1981)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: An inlet technology development program for future supersonic cruise aircraft is in progress. Areas being emphasized are inlet aerodynamic and control system design requirements for efficient and reliable operation. Off-design conditions, such as angle of incidence, starting, and noise abatement are major considerations. Flow analysis procedures are being developed to predict the internal inviscid and viscous flows in axisymmtric supersonic inlets for these operating conditions. Also under development are control systems that will have significant interfaces with the engine control system, the flight control system, and the airframe avionics system. The analytical methods are being supported and validated with representative experiments.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-1598 , Joint Propulsion Conference; Jul 27, 1981 - Jul 29, 1981; Colorado Springs, CO; US
    Format: text
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    NASA TECHNICAL REPORTS
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