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NASA NTRS2006

Boundary-Layer-Ingesting Inlet Flow Control

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 (DPCPavg) 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.

Owens, Lewis R.↗
NASA NTRS2006

Boundary-Layer-Ingesting Inlet Flow Control

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(sub 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.

Owens, Lewis R.↗
NASA NTRS1975

Effectiveness of spoilers on the GA(W)-1 airfoil with a high performance Fowler flap

Two-dimensional wind-tunnel tests were conducted to determine effectiveness of spoilers applied to the GA(W)-1 airfoil. Tests of several spoiler configurations show adequate control effectiveness with flap nested. It is found that providing a vent path allowing lower surface air to escape to the upper surface as the spoiler opens alleviates control reversal and hysteresis tendencies. Spoiler cross-sectional shape variations generally have a modest influence on control characteristics. A series of comparative tests of vortex generators applied to the (GA-W)-1 airfoil show that triangular planform vortex generators are superior to square planform vortex generators of the same span.

Wentz, W. H., Jr.↗
NASA NTRS1970

Aircraft vortex wakes

Aircraft generated vortex wakes and core air motions hazards for encountering light airplane

Boswinkle, R. W., Jr.↗
NASA NTRSDate not supplied

CFD Simulation of an Inward-Turning Supersonic Inlet Unstart at Flight Mach Number 1.7

Interest in inward-turning supersonic inlets renewed during NASA’s Commercial Supersonic Technology Project, as a means of reducing disturbances caused by propulsion integration. Inward-turning designs have the potential to greatly reduce the strength of shock waves emanating outward from the nacelle, and therefore their contribution to the overall sonic boom signature. For top-mounted, or over-wing installations, disturbances to aircraft aerodynamics are reduced. The nature of the inward-turning geometry results in a boundary-layer thickening effect, opposite that of the familiar thinning of an outward-turning cone boundary-layer. The application of a small amount of boundary-layer bleed at the throat has been shown to alleviate the negative effects of the boundary-layer on subsonic diffuser performance. Boundary-layer bleed, however, comes at the expense of increased complexity, and external disturbances at bleed exits that would partially defeat the purpose of the inward-turning scheme. Vortex generators were proposed as an alternative to bleed, and an initial CFD study of their effectiveness in this application forms the basis of the present work. A source-term model in the Wind-US CFD code was used to model the effects of an array of vortex generators placed at the entrance to the subsonic diffuser. Baseline results with no vortex generators compared favorably to previously reported results with no bleed at back-pressures near the design point, but stability margin in the present case was limited by a curious non-linear phenomena similar to unstart in mixed-compression inlets. While some improvement was noted with vortex generators, their effectiveness was diminished by the limited stability margin. A time-accurate calculation was subsequently performed to better understand this process. Results of this calculation are presented and instantaneous solutions compared with steady-state, subcritical results previously reported. The stability margin limitation is attributed to the use of a constant-pressure outflow boundary condition, which differed from that used in the prior simulations.

Pre-Compressor Cooling, Water Injection, Hypersoni↗
NASA NTRS1992

Application of Computational Fluid Dynamics to the Study of Vortex Flow Control for the Management of Inlet Distortion

The present study demonstrates that the Reduced Navier-Stokes code RNS3D can be used very effectively to develop a vortex generator installation for the purpose of minimizing the engine face circumferential distortion by controlling the development of secondary flow. The computing times required are small enough that studies such as this are feasible within an analysis-design environment with all its constraints of time and costs. This research study also established the nature of the performance improvements that can be realized with vortex flow control, and suggests a set of aerodynamic properties (called observations) that can be used to arrive at a successful vortex generator installation design. The ultimate aim of this research is to manage inlet distortion by controlling secondary flow through an arrangements of vortex generators configurations tailored to the specific aerodynamic characteristics of the inlet duct. This study also indicated that scaling between flight and typical wind tunnel test conditions is possible only within a very narrow range of generator configurations close to an optimum installation. This paper also suggests a possible law that can be used to scale generator blade height for experimental testing, but further research in this area is needed before it can be effectively applied to practical problems. Lastly, this study indicated that vortex generator installation design for inlet ducts is more complex than simply satisfying the requirement of attached flow, it must satisfy the requirement of minimum engine face distortion.

Anderson, Bernhard H.↗
NASA NTRS1960

Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed

A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.

Smith, Willard G.↗
NASA NTRS1973

Wind tunnel simulation of full scale vortices

An experimental investigation has been conducted to determine the important scaling parameters for the flow in the core region of a vortex generated by a rectangular wing tip. The effect of an unconventional planform, the ogee tip, on the tip vortex is also determined. For rectangular planform wings, the measured vortex core diameter to chord ratios, peak tangential velocity ratios, and axial velocity ratios are shown to be functions only of wing lift coefficient and elapsed time from vortex formation, and appear to be independent of both Mach number and Reynolds number. The peak tangential velocities in the diffuse vortex generated by the ogee tip are only 25 percent of those in the vortex generated by the rectangular wing.

Rorke, J. B.↗
NASA NTRS2012

Calibration of the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel (1996 and 1997 Tests)

There were several physical and operational changes made to the NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel during the period of 1992 through 1996. Following each of these changes, a facility calibration was conducted to provide the required information to support the research test programs. Due to several factors (facility research test schedule, facility downtime and continued facility upgrades), a full test section calibration was not conducted until 1996. This calibration test incorporated all test section configurations and covered the existing operating range of the facility. However, near the end of that test entry, two of the vortex generators mounted on the compressor exit tailcone failed causing minor damage to the honeycomb flow straightener. The vortex generators were removed from the facility and calibration testing was terminated. A follow-up test entry was conducted in 1997 in order to fully calibrate the facility without the effects of the vortex generators and to provide a complete calibration of the newly expanded low speed operating range. During the 1997 tunnel entry, all planned test points required for a complete test section calibration were obtained. This data set included detailed in-plane and axial flow field distributions for use in quantifying the test section flow quality.

Arrington, E. Allen↗
NASA NTRS1997

Development of a Low-Reynolds Number, Nonlinear kappa-epsilon Model for the Reduced Navier-Stokes Equations

Previous work at NASA LeRC has shown that flow distortions in aircraft engine inlet ducts can be significantly reduced by mounting vortex generators, or small wing sections, on the inside surface of the engine inlet. The placement of the vortex generators is an important factor in obtaining the optimal effect over a wide operating envelope. In this regard, the only alternative to a long and expensive test program which would search out this optimal configuration is a good prediction procedure which could narrow the field of search. Such a procedure has been developed in collaboration with NASA LeRC, and results obtained by NASA personnel indicate that it shows considerable promise for predicting the viscous turbulent flow in engine inlet ducts in the presence of vortex generators. The prediction tool is a computer code which numerically solves the reduced Navier-Stokes equations and so is commonly referred to as RNS3D. Obvious deficiencies in RNS3D have been addressed in previous work. Primarily, it is known that the predictions of the mean velocity field of a turbulent boundary layer flow approaching separation are not in good agreement with data. It was suggested that the use of an algebraic mixing-length turbulence model in RNS3D is at least partly to blame for this. Additionally, the current turbulence model includes an assumption of isotropy which will ultimately fail to capture turbulence-driven secondary flow known to exist in noncircular ducts.

Boger, David A.↗
NASA NTRS2010

Boundary-layer-ingesting inlet flow control system

A system for reducing distortion at the aerodynamic interface plane of a boundary-layer-ingesting inlet using a combination of active and passive flow control devices is disclosed. Active flow control jets and vortex generating vanes are used in combination to reduce distortion across a range of inlet operating conditions. Together, the vortex generating vanes can reduce most of the inlet distortion and the active flow control jets can be used at a significantly reduced control jet mass flow rate to make sure the inlet distortion stays low as the inlet mass flow rate varies. Overall inlet distortion, measured and described as average SAE circumferential distortion descriptor, was maintained at a value of 0.02 or less. Advantageous arrangements and orientations of the active flow control jets and the vortex generating vanes were developed using computational fluid dynamics simulations and wind tunnel experimentations.

Owens, Lewis R.↗
NASA NTRS1992

Separation control on high Reynolds number multi-element airfoils

Small surface-mounted vortex generators were investigated as means for the control of a boundary-layer separation on a 2D single-flap three-element high-lift system at near-flight Reynolds numbers and in landing configurations. Wind-tunnel results obtained for small vane-type vortex generators mounted on a multielement airfoil showed that vortex generators as small as 0.18 percent of total chord can effectively reduce or eliminate boundary-layer separation on the flap at approach conditions. It was found that both the outerrotating and the corotating streamwise vortices were effective in reducing flow separation.

Lin, John C.↗
NASA NTRS1954

Investigation of Conical Subsonic Diffusers for Ram-jet Engines

The efficiency of a 30 degree conical diffuser was improved as much as 20 percent and separation was eliminated by the use of vortex generators. The use of splitter cones gave only small efficiency gains, but relatively uniform diffuser-outlet velocity profiles were obtained with the better designs. A configuration which incorporated both vortex generators and a splitter cone gave efficiencies higher than those obtained with any other splitter-cone configuration and also higher than those obtained when the same vortex generators were used without the splitter cone. Moderate diffuser-efficiency increases were obtained with guide vanes; however, in all cases combustion occurred in the vane wakes.

Farley, John M↗
NASA NTRS1972

Microfog lubricant application system for advanced turbine engine components, phase 2. Tasks 3, 4 and 5: Wettability and heat transfer of microfog jets impinging on a heated rotating disc, and evaluation of reclassifying nozzles and a vortex mist generator

The wettabilities and heat transfer rates of microfog jets (oil-mist nozzle flows) impinging on a heated rotating disc were determined under an inert atmosphere of nitrogen at temperatures ranging from 600 to 800 F. The results are discussed in relation to the various factors involved in the microfog lubricant application systems. Two novel reclassifying nozzles and a vortex mist generator were also studied.

Shim, J.↗
NASA NTRS1981

Vortex-flow aerodynamics - An emerging design capability

Promising current theoretical and simulational developments in the field of leading edge vortex-generating delta, arrow ogival wings are reported, along with the history of theory and experiment leading to them. The effects of wing slenderness, leading edge nose radius, Mach number and incidence variations, and planform on the onset of vortex generation and redistribution of aerodynamic loads are considered. The range of design possibilities in this field are consequential for the future development of strategic aircraft, supersonic transports and commercial cargo aircraft which will possess low-speed, high-lift capability by virtue of leading edge vortex generation and control without recourse to heavy and expensive leading edge high-lift devices and compound airfoils. Attention is given to interactive graphics simulation devices recently developed.

Campbell, J. F.↗
NASA NTRS1996

Experimental Investigation of the Performance of a Mach-2.7 Two-dimensional Bifurcated Duct Inlet with 30 Percent Internal Contraction

An experimental study was conducted to determine the performance of a two-dimensional, mixed-compression bifurcated duct inlet system designed for a free-stream Mach number of 2.7. Thirty percent of the supersonic area contraction occurred internally. A movable ramp was used to vary the contraction ratio for off-design operation. Boundary layer bleed regions were located on the cowl, centerbody, and sidewall surfaces. There were also provisions for vortex generators on the cowl and centerbody of the subsonic diffuser. Data were obtained over the Mach number range of 2.0 to 2.8 and at angles of yaw from 0 deg. to the maximum value prior to inlet un-start. The test at Mach 2.8 was to obtain data for an over- speed condition. The Reynolds number varied from 2.5 to 2.3 million/ft for Mach numbers above 2.5. At Mach numbers of 2.5 and lower, the Reynolds number was set at 2.5 million/ft. Bleed patterns, vortex generator patterns, and ramp position were varied, and three inlet configurations were selected for more extensive study. Two of these configurations had self-starting capability. The self-starting configuration that was developed produced 89 percent total pressure recovery at the compressor face station with 6.8 percent total bleed. The compressor face distortion was about 16 percent. Vortex generators were extremely effective in re-distributing flow but were not as effective in reducing distortion. Excellent flow symmetry was achieved between the separated halves of the inlet, and twin-duct instability was not observed. The ramp tip shock was steeper than expected. This caused the cowl lip shock to be reflected from the ramp instead of being cancelled at the shoulder. However, peak recovery at the throat was still obtained with the ramp near the design position.

Wasserbauer, J. F.↗
NASA NTRS1995

Experimental Study of Lift-Enhancing Tabs on a Two-Element Airfoil

The results of a wind-tunnel test are presented for a two-dimensional NASA 63(sub 2)-215 Mod B airfoil with a 30% chord single-slotted flap. The use of lift-enhancing tabs (similar to Gurney flaps) on the lower surface near the trailing edge of both elements was investigated on four nap configurations. A combination of vortex generators on the flap and lift-enhancing tabs was also investigated. Measurements of surface-pressure distributions and wake profiles were used to determine the aerodynamic performance of each configuration. By reducing flow separation on the flap, a lift-enhancing tab at the main-element trailing edge increased the maximum lift by 10.3% for the 42-deg flap case. The tab had a lesser effect at a moderate flap deflection (32 deg) and adversely affected the performance at the smallest flap deflection (22 deg). A tab located near the flap trailing edge produced an additional lift increment for all flap deflections. The application of vortex generators to the flap eliminated lift-curve hysteresis and reduced flow separation on two configurations with large flap deflections (greater than 40 deg). A maximum-lift coefficient of 3.32 (17% above the optimum baseline) was achieved with the combination of lift-enhancing tabs on both elements and vortex generators on the flap.

Storms, Bruce L.↗