This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP avg ) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow. Nomenclature a = distance between the inlet lip highlight station and the inlet throat station (see fig. 3), in. A C = inlet capture (highlight) area; area enclosed by inlet highlight (see fig. 3 for highlight definition) and tunnel wall, in. = distance between the inlet highlight height and the inlet throat height (see fig. 3), in. D 2 = duct diameter at AIP (see fig. 3), in. DPCP avg = average SAE circumferential distortion descriptor DPRP i = SAE radial distortion descriptor for ring i on AIP total-pressure rake h = height of vortex generator, in. H = boundary-layer shape factor, δ * /θ H i = height of inlet throat (see fig. 3), in. H max = maximum height of inlet cowl (see fig. 3), in. ΔH = distance between inlet throat centroid and duct exit centroid (see fig. 3), in. i = ring number on AIP total-pressure rake, value increases from 1 in hub region to 5 in tip region L = length of inlet duct from throat to AIP (see fig. 3 https://ntrs.nasa.gov/search.jsp?R=20060004990 2018-05-12T20:30:27+00:00ZAmerican Institute of Aeronautics and Astronautics 2 P t = free-stream total pressure, psi P t,2,avg = area weighted average total pressure at AIP P t,2,avg /P t = inlet recovery pressure ratio Rn/ft = R...
A current NASA Research Announcement (NRA) project being conducted by Georgia Tech Research Institute (GTRI) personnel and NASA collaborators includes the development of Circulation Control (CC) blown airfoils to improve subsonic aircraft high-lift and cruise performance. The emphasis of this program is the development of CC active flow control concepts for both high-lift augmentation, drag control, and cruise efficiency. A collaboration in this project includes work by NASA research engineers, whereas CFD validation and flow physics experimental research are part of NASA's systematic approach to developing design and optimization tools for CC applications to fixed-wing aircraft. The design space for CESTOL type aircraft is focusing on geometries that depend on advanced flow control technologies that include Circulation Control aerodynamics. The ability to consistently predict advanced aircraft performance requires improvements in design tools to include these advanced concepts. Validation of these tools will be based on experimental methods applied to complex flows that go beyond conventional aircraft modeling techniques. This paper focuses on recent/ongoing benchmark high-lift experiments and CFD efforts intended to provide 2-D CFD validation data sets related to NASA's Cruise Efficient Short Take Off and Landing (CESTOL) study. Both the experimental data and related CFD predictions are discussed.
A new high Reynolds number test capability for boundary layer ingesting inlets has been developed for the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel.Using this new capability, an experimental investigation of four S-duct inlet configurations was conducted. A computational study of one of the inlets was also conducted using a Navier-Stokes solver.The objectives of this investigation were to 1) develop a new high Reynolds number inlet test capability for flush-mounted inlets, 2) provide a database for CFD tool validation, 3) evaluate the performance of S-duct inlets with large amounts of boundary layer ingestion and 4) provide a baseline inlet for future inlet flow-control studies. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on duct exit diameter) from 5.1 million to a full-scale value of 13.9 million, and inlet mass-flow ratios from 0.39 to 1.58 depending on Mach number. Results of the experimental study indicate that inlet pressure recovery generally decreased and inlet distortion generally increased with increasing Mach number. Except at low Mach numbers, increasing inlet mass-flow increased pressure recovery and increased distortion. Increasing the amount of boundary layer ingestion or ingesting a boundary layer with a distorted profile decreased pressure recovery and increased distortion. Finally, increasing Reynolds number had almost no effect on inlet distortion but increased inlet recovery by about one-half percent at a Mach number near cruise. The computational results captured the inlet pressure recovery and distortion trends with Mach number and inlet mass-flow well; the reversal of the pressure recovery trend with increasing inlet mass-flow at low and high Mach numbers was predicted by CFD. However, CFD results were generally more pessimistic (larger losses) than measured experimentally.
Boundary layer ingestion (BLI) is explored as a means to improve overall system performance for a Blended Wing Body configuration. The benefits of BLI for vehicle system performance benefit are assessed with a process derived from first principles suitable for highly-integrated propulsion systems. This performance evaluation process provides a framework within which to assess the benefits of an integrated BLI inlet and lays the groundwork for higher-fidelity systems studies. The results of the system study show that BLI provides a significant improvement in vehicle performance if the inlet distortion can be controlled, thus encouraging the pursuit of active flow control (AFC) as a BLIenabling technology.The effectiveness of active flow control in reducing engine inlet distortion was assessed using a 6% scale model of a 30% BLI offset, diffusing inlet. The experiment was conducted in the NASA Langley Basic Aerodynamics Research Tunnel with a model inlet designed specifically for this type of testing. High mass flow pulsing actuators provided the active flow control. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion was determined by 120 total pressure measurements located at the aerodynamic interface plane. The test matrix was limited to a maximum freestream Mach number of 0.15 with scaled mass flows through the inlet for that condition. The data show that the pulsed actuation can reduce distortion from 29% to 4.6% as measured by the circumferential distortion descriptor DC60 using less than 1% of inlet mass flow. Closed loop control of the actuation was also demonstrated using a sidewall surface static pressure as the response sensor.
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