Results of numerical simulations of active flow control applied to an airfoil using the Reynoldsaveraged Navier-Stokes equations are presented. The simulations are first compared with the poststall separation control experiments of Seifert et al.1'12 on a NACA0015 at 1.2 x 106 chord Reynolds number. The jet is introduced tangential to the surface at the leading edge of the airfoil. The calculated lift increments are in good agreement with the experimental data. Two flow-control techniques for a NACA0012 airfoil at a chord Reynolds number of 8.5 x 10 6 are investigated. The first technique utilizes a small, 0.5% chord, steady jet, and the second method employs a synthetic jet of a similar scale. Performance benefits are obtained by placing the actuators very near the airfoil leading edge on the suction surface. A significant increase in lift (29%) is obtained using the synthetic jet actuator in the post-stall regime. At lower lift, the steady jet actuator significantly reduces drag by rotating the lift vector upstream.

Active flow control applied to an airfoil

A Spalart-Allmaras turbulence model implementation in a discontinuous Galerkin solver for incompressible flows

This paper investigates the effects using high lift zero-net mass-flux Co-Flow Jet (CFJ) active flow control airfoil for aircraft control surfaces with plain flaps and with no flap. The goal is to reduce the size and weight of conventional aircraft control surfaces and save energy expenditure. Two-dimensional simulation of NACA 0012 airfoil used as a control surface is conducted for parametric trade study using a Reynolds-averaged Navier-Stokes (RANS) solver with Spalart-Allmaras (SA) model. A 5th order WENO scheme for the inviscid flux and a 4th order central differencing for the viscous terms are used to resolve the Navier-Stokes equations. The 2D numerical studies indicate that the CFJ airfoil for aircraft control surfaces with a plain flap can dramatically increase the lift coefficient and aerodynamic efficiency simultaneously compared with the conventional control surface with the same size of flap and deflection angle. CFJ airfoil control surface shows great potential to substantially reduce the size and weight of conventional aircraft control surfaces with high control authority. A series of trade study is done based on NACA0012 airfoil for control surface. The CFJ airfoil is modified from the baseline NACA0012 airfoil by translating the upper surface downward by 0.1%C. A constant deflection angle of 30° is used. The final preferred configuration has the flap length of 35%C, deflection angle of 30°, injection location at 2%C from leading edge, injection slot size of 0.5%C, and suction slot right upstream of the flap with the size twice larger than the injection size. The final trade study is to investigate the effect of injection jet momentum coefficient Cμ at 0.05, 0.15, 0.25. The lower Cμ value of 0.05 is the most energy cost effective to increase the lift coefficient. Comparing with the baseline airfoil with the same flap size and deflection angle, at sideslip angle of 0°, the case of Cμ=0.05 of the final configuration achieves a lift coefficient increase by 106.4% from CL=1.09 to 2.25 at very low power coefficient of 0.0285. At the same time, it substantially reduces the drag by 67.17%. All these compound effects result in an increase of aerodynamic efficiency(including CFJ power consumption) by 232.2%. In other words, while the CFJ control surface substantially increases the lift, it simultaneously reduces the net energy cost in a dramatical manner. This even does not count the additional benefit due to the reduced control surface size and weight Finally, CFJ airfoil with no flap is also simulated at injection jet momentum coefficient Cμ=0.05, 0.10, 0.15, 0.20, 0.25 and 0.30. The result shows that the maximum lift coefficients of 3.048 (an increase of 114%) is achieved at Cμ=0.30 with a reduced drag. The aerodynamic efficiency of the flapless control surface is not studied in this work and will be reported in future. The results indicate that flapless control surface may be a feasible option. 1 Ph.D. Candidate of Nanjing University of Aeronautics and Astronautics, visiting Student at University of Miami 2 Ph.D. Candidate 3 Ph.D. Candidate, AIAA student member 4 Postdoc Researcher, Ph.D., AIAA member 5 Ph.D. Candidate 6 Professor, ASME Fellow, AIAA associate Fellow D ow nl oa de d by G ec he ng Z ha o n Ju ly 9 , 2 01 8 | h ttp :// ar c. ai aa .o rg | D O I: 1 0. 25 14 /6 .2 01 830 67 2018 Applied Aerodynamics Conference June 25-29, 2018, Atlanta, Georgia 10.2514/6.2018-3067 Copyright © 2018 by all the authors of this paper . Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. AIAA AVIATION Forum

Aircraft control surfaces using co-flow jet active flow control airfoil

This paper numerically investigates drag minimization techniques of the CFJ control surface when it is not activated at cruise condition. Three methods are studied to minimize the control surface drag. One is to let the CFJ produce a very light jet at the airfoil surface to decrease the drag at a low power. This method is effective at various cruise Mach numbers. Compared with the baseline, it decreases the corrected drag coefficient ((CD)c) by 11.4%, 13.0% and 14.6% at Ma∞=0.15, 0.45 and 0.7. The second method is to cover the injection and suction slot with a small movable surface segment and turn the CFJ off at cruise. This brings the CD of CFJ airfoil to the same level as the baseline airfoil. The third method is to simply turn the micro-compressor actuator off and leave the CFJ ducts open with no jet. This method does increase the drag. As a result of the increased lift coefficient that allows reducing the control surface area, the overall drag of the control surface will be significantly decreased to benefit the aircraft system.

Drag Minimization of Co-Flow Jet Control Surfaces at Cruise Conditions

The analysis of two dimensional (2D) flow over NACA 0012 airfoil is validated with NASA Langley Research Center validation cases. The k-ω shear stress transport (SST) model is utilized to predict the flow accurately along with turbulence intensities 1% and 5% at velocity inlet and pressure outlet respectively. The computational domain is composed of 120000 structured cells. In order to enclose the boundary layer method the enhancement of the grid near the airfoil is taken care off. This validated simulation technique is further used to analyse aerodynamic characteristics of plain flapped NACA 0012 airfoil subjected to different flap angles and Mach number. The calculation of lift coefficients (CL), drag coefficients (CD) and CL/CD ratio at different operating conditions show that with increasing Mach number (M) CL increases but CD remains somewhat constant. Moreover, a rapid drastic decrease is observed for CL and an abrupt upsurge is observed for Cd with velocity approaching to the sonic velocity. In all cases range and endurance are decreased, as both values of CL/CD and √CL/CD are declined.

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Computational Study of Flow around a NACA 0012 Wing Flapped at Different Flap Angles with Varying Mach Numbers

Tests were conducted in the 2-D test section of the Langley 0.3-meter Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th power. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Tabulated pressure distributions and integrated force and moment coefficients are presented as well as plots of the surface pressure distributions. The data are presented uncorrected for wall interference effects and without analysis.

Pressure distributions from high Reynolds number transonic tests of an NACA 0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

An internal strain-gage balance designed and constructed in Europe for use in cryogenic wind tunnels has been tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. Part of the evaluation was made at equilibrium balance temperatures and it consisted of comparing the data taken at a tunnel stagnation temperature of 300 K with the data taken at 200 K and 110 K while maintaining either the Reynolds number or the stagnation pressure. A sharp-leading-edge delta-wing model was used to provide the aerodynamic loading for these tests. Results obtained with the balance during the force tests were found to be accurate and repeatable both with and without the use of a convection shield on the balance. An additional part of this investigation involved obtaining data on the transient temperature response of the balance during both normal and rapid changes in the tunnel stagnation temperature. The variation of the temperature with time was measured at three locations on the balance near the physical locations of the strain gages. The use of a convection shield significantly increased the time required for the balance to stabilize at a new temperature during the temperature response tests.

Aerodynamic measurements and thermal tests of a strain-gage balance in a cryogenic wind tunnel

A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

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Acquilla S. Hill

Papers

Pressure distributions from high Reynolds number transonic tests of an NACA 0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

Aerodynamic measurements and thermal tests of a strain-gage balance in a cryogenic wind tunnel

High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel