Document
F_
AIAA 2002-3157
An Active Flow Circulation
Controlled Flap Concept for General
Aviation Aircraft Applications
Gregory S. Jones, Sally A. Viken, Anthony E. Washburn, Luther N. Jenkins and C. Mark Cagle.
NASA Langley Research Center
Hampton, VA
1st AIAA Flow Control Conference
June 24-26, 2002/St. Louis, Missouri
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1801 Alexander Bell Drive, Suite 500, Reston, VA 20191--4344
An Active Flow Circulation Controlled Flap Concept for
General Aviation Aircraft Applications
G. S. Jones* S. A. Viken* A.E. Washburn* L. N. Jenkins* C.M. Cagle** Flow Physics & Control Branch NASA Langley Research Center Abstract A recent focus on revolutionary aerodynamic concepts has highlighted the technology needs of general aviation and personal aircraft. New and stringent restrictions on these types of aircraft have placed high demands on aerodynamic performance, noise, and environmental issues. Improved high lift performance of these aircraft can lead to slower takeoff and landing speeds that can be related to reduced noise and crash survivability issues.
Circulation Control technologies have been around for 65 years, yet have been avoided due to trade offs of mass flow, pitching moment, perceived noise etc. The need to improve the circulation control technology for general aviation and personal air-vehicle applications is the focus of this paper. This report will describe the development of a 2-D General Aviation Circulation Control (GACC) wing concept that utilizes a pulsed pneumatic flap.
Subscripts: Symbols Free stream Conditions A Area (_) J Jet at slot exit b Span (inches) o Stagnation Condition CL Uft Coefficient BAL Measurements w/strain gage balance Co Drag Coefficient EQUIV Equivalent (referenced to drag) Ca Moment Coefficient C,, Momentum Coefficient Introduction C Chord (inches) In recent years, there has been an increasing CCW Circulation Controlled Wing interest in revolutionary concepts applied to general D Drag (Ibs) aviation and personal aircraft I shown in figure 1. New DC Duty Cycle (Time On/Time Off) and stringent requirements on these types of aircraft E Mean Voltage include aerodynamic performance, noise, and e' Fluctuating Voltage environmental issues. The use of Pulsed Pneumatic h Slot height (inches) High Lift Technology has the potential to revolutionize LE Leading Edge aircraft systems by reducing wing area, reducing part L Lift (Ibs) count, lowering weight, and reducing potential runway M Pitching Moment (in-lbs) take off and landing requirements. This paper will give a m mass flow (Ib/sec) brief background and review of circulation control P Pressure (Ib/in 2 or Ib/tt 2) physics then describe the development of a 2-D General p' FluctUating Pressure (Ib/in 2 or Ib/ft 2) Aviation Circulation Control (GACC) pulsed wing r Trailing edge radius (inches) concept that utilized CFD and wind tunnel experiments.
U Velocity (ft/sec) u' Fluctuating Velocity (ft/sec) q Dynamic Pressure (Ib/ft 2) S Wing plan form area (_) SCFM Standard Mass Flow (ft3/min) (Expanded to 14.7 psia & 72°F) SPL Sound Pressure Level (dB) TE Trailing Edge T Static Temperature (°R) w Slot Width (inches) Angle of attack (degrees) p density (Lbm/ft 3) r Circulation • RESEARCH ENGINEER *" DESIGN ENGINEER Figure 1. Artist concept of a personal air vehicle Copy_ight(_2002 by the American Instit_e of Aeronautics and _utlcs, Inc.
NO oopyrlght Is asserted in the United States under Title 17, U.S, Code, The U.S Governtrmnt has 8 royalty-free license to e_erctle i_11 rights ufKler 1tie copyright claimed herein for Govemmenlat Purposes All other tights are resewed by the AIAA 2002-3157 copyr_hl OWrte=" Once the jet separates from the Coanda surface Background it penetrates the flow field resulting in a large deflection The definition of "Circulation Control" is of the streamlines producing a pneumatic camber similar to a mechanical high lift system, figure 3. The strictly related to the circulation characteristics combination of the Coanda separation and the jet around any aerodynamic body and can be penetration will move the rear stagnation point forward controlled or managed with many different control on the lower surface and move the leading edge schemes including airfoil shape and shape change, stagnation point aft on the lower surface. As the jet flaps, ailerons, blowing, suction, etc. Traditionally velocity is increased, these stagnation points move Circulation Control Wings (CCW) are restricted to a toward each other resulting in more circulation. Under pneumatic modification of the flow field through a ideal circumstances (e.g. circular cylinder) the leading Coanda effect. 2 This well known effect was named edge stagnation point and trailing edge stagnation point after Henri Coanda who accidentally discovered it in will form a singularity. If the thrust effects from the jet Paris in 1935. Coanda was trying to deflect the were ignored, this would correspond to a lift coefficient exhaust for an engine only to entrain the hot gas limit of 4=.
that resulted in destroying the aircraft. This Coanda effect shown in figure 2 can be described by a 2-D wall bounded jet that exits from a slot tangential to a convex curved surface. The wall bounded jet flows along the surface and has the nature of a boundary layer near the wall but becomes that of a free jet at a larger distance from the wall. 3 The degree of jet turning can be related to the slot height, surface radius, jet velocity, and the Coanda surface Figure 3 Coanda influence on streamlines / COANDA __ ...... . TURNING '1 The aerodynamic characteristics of Circulation .............,,_ ANGLE(_) !
Control Wings (CCW) have been experimentally and numerically studied for more than 65 years, s's'7 Many of
/
these studies have concentrated on trailing edge shape, slot height, and blowing rates. It is obvious that the airfoil shape plays a major role in performance. Various studies utilized airfoil geometries for specific applications such as helicopter rotors, wings, sails, and airfoils with and without camber, etc. Results of these studies highlighted geometric features that affect airfoil Figure 2 Trailing edge example of Coanda effect performance such as the ratio of the trailing edge radius to chord (r/C), slot height to chord ratio (h/C), slot height and in some cases exceed 180 ° . The jet will remain to radius ratio (h/r), and Coanda surface shape. In attached to the curved surface because of a general, the larger the trailing edge radius, the more balance between the sub-ambient pressure in the effective the Coanda effect has on lift due to increased jet sheet and the centrifugal force around the surface area. This is good for the high lift configuration curvature of the surface. Although the Coanda but once the airfoil has reached cruise conditions, there effect is very effective for boundary layer control is large drag penalty due to the blunt trailing edge.
(BLC), the interest in this technology comes from its It has been found that the use of steady jets, ability to further augment the circulation and lift with even at very small mass flow rates, yielded lift flow turning and control of leading edge streamlines, coefficients that are comparable or superior to and thus the name Circulation Control (CC).
conventional high-lift systems. Several CCW high lift Lanchester, Kutta, and Joukowski laid the studies 8 have shown lift results that approach sectional foundation for a quantitative theory relating the lift to lift coefficients of 9. This can be compared to traditional an infinite wing through the integration of the mechanical high lift systems that approach maximum lift velocity field along a streamline. 4 coefficients of 6 and require complex mechanical [" = _V.dL systems shown in figure 4. These mechanical systems Where lift is go beyond the scope of most general aviation aircraft. A Lift = p.U. F more reasonable general aviation high lift system would AtAA 1 = Flow Control Conference 2 AIAA 2002-3157 both experiments were limited in scope and little was be either a simple hinged flap or a single slotted revealed about the physics of the phenomena.
Fowler flap. The lift performance of a typical high lift system is shown in figure 5. 9 Since CCW systems General Aviation Circulation Control offer such a large potential gain in lift performance, the application to STOL aircraft seems appropriate. The GACC test program is intended to address L.E. DROOP HINGED FLAP technology issues, such as scaling, mass flow, and noise requirements. A 2-D flow physics supercritical airfoil model TM (figure 6) with dual slotted circulation control capability has been designed and built for low L.E. SLAT 1-SLOT FLAP speed testing in the LaRC Basic Aerodynamic Research Tunnel (BART). The primary objective of the program is FLAT KRUEGER 2-SLOT FLAP to evaluate the benefits of pulsed circulation control and to reduce the mass flow requirements for a given lift performance as well as reduce the cruise drag penalty associated with a large circulation control trailing edge.
CURVED KRUEGER 3-SLOT FLAP UPPER STEADY ACTUATOR UPPER DIFFUSER SLOT MANIFOLD\ X k _ Figure 4. Conventional High Lift Components that a
-4(------_- 9.40' \ --'_
CCW high lift system could potentially replace 4.5 .... i .... I .... i .... , .... " ' ''_lP .... ' 4 EXPERIMENTAL "_ LOWER STEADY \ PULSED / MANIFOLD \ ACTUATOR LOW/ U.I 35 ACTUATOR SLOT MANIFOLD Ii W Figure 6. 2-Dimensional 17% Supercritical General
_z /
Aviation Circulation Controlled Airfoil with a circular _25 trailing edge r/C: 2% .J 2 SOLID SYMBOLS: CFD The optimization of high lift and cruise
- /7
performance with one airfoil shape gives rise to the pneumatic flap concept 15'16 This concept is based on the 15 .... = .... = .... I .... i .... _ .... I ....
ability to switch from a high lift configuration to a cruise
-s 0 5 10 15 20 25 30
configuration without utilizing any mechanical systems.
AOA (DEGREES) Having two independent blowing systems allows one to Figure 5. Comparison of computed (FUN2D) and have such a multi-function system that can be used for experimental results for three-element airfoil at high lift systems and flight control systems such as Mach of 0.20 and Re of 9.x 108(Ref 9) ailerons and air brakes. Moving from a high lift to a cruise configuration is dependent on the upper and Through the 1960's and 1980's the U.S.
lower blowing ratios and the free stream velocity.
Navy evaluated numerous CCW concepts. 1° To As the Coanda effects are modifying the flow- summarize many of these studies and demonstrate field at the trailing edge, the leading edge stagnation the bleed affects associated with CCW, a simultaneously moves downstream. This imposes a Circulation Control wing was flight-tested on an A-6 large pressure gradient at the leading edge that can aircraft. This demonstrator showed significant lead to premature boundary layer separation and airfoil improvements in takeoff and landing characteristics.
stall. To avoid conventional leading edge slats or other The maximum lift coefficient went from an un-blown flow control techniques a blunt leading edge is desired.
to blown case of 2.1 to 3.9 respectively. The A 17% supercritical airfoil shape based on the GAW(1) approach speeds decreased from 118 to 76 knots 11.
was modified at the trailing edge to create the GACC During the mid-1970's two efforts were profile. In addition to providing a blunt leading edge completed that focused on pulsed blowing this airfoil also provides a sufficient internal volume to associated with circulation control? 2' 13 Results from house the three pressure manifolds and pulsed these experiments indicated that pulsed blowing actuator.
reduced the mass requirements for CCW. However AIAA 1= Flow Control Conference 3 AIAA 2002-3157 The GACC model was tested for high lift highlights the non-linearity of the momentum coefficient and cruise configurations by modifying the mass due to the density ratio.
flow through the upper and/or lower slots. Each slot Using the momentum formulation described flow was independently controlled and will be above does not accurately characterize the physics discussed later in the text.
related to airfoil performance for different slot heights.
Initial testing focused on lift characteristics Nominally smaller slot heights yield a larger return in lift consistent with general aviation aircraft. A target lift coefficient at constant C_t than do larger slot heights.
coefficient of 3 was determined to be adequate or An empirical technique 6 described by equation 5 is often exceed most general aviation and personal air used to collapse the performance data.
vehicle requirements. To determine the experimental test matrix and instrumentation requirements, a 2-D CFD effort was used to quantify flow parameters such as boundary layer separation, Z$O - slot velocity profiles, mass flow, model surface pressure profiles, internal plenum pressures, lift,
,,o
drag, and pitching moment.
The two most popular Coanda shapes typically used for CCW applications are based on circular and elliptic profiles. For low speed applications such as high lift, the circular profile is more effective than the elliptic shape. However the i '°° /A___._ _ ......
elliptic shape is more effective at high-speed cruise conditions than the circular shape.
A 2% r/C circular trailing edge shape was chosen as a baseline for the GACC model. The o.o zs.o so.o 7s.o 1 ooo 1 zs.o trade off of high lift performance and cruise drag is (ujet/uo) _ based on optimizing the high lift system first then Figure 7. Mass flow requirements for a circulation utilizing the dual blowing to optimize the cruise drag.
control wing at a q: 10 psf and To: 75°F.
Theoretical Conslderatlonl_ Pulsed theoretical considerations The momentum coefficient C_ is a critical For the time dependent pulsed flows the jet velocity will parameter in understanding the efficiencies of blown be divided into a mean and fluctuating component.
-- r systems such as the GACC and is defined as Uj=Uj+uj (Eq.6) Substituting into equation 4 the total momentum C tJ Thrust ,;,( Uj ) coefficient becomes: qS q(C)(b) (Eq. 1) C_=C_+C_ (Eq.7) where where m = pjUjAj (Eq.2) and C.' = .-[ (C-_j_=--=)(Eq.8) and
r flrP=
U j= 2yR(TDUCT)
(Eq.9)
y-1 I -_PD--_-u_) ) (Eq.3)
'
It should be noted that the jet velocity has been Equation 8 assumes that the 2 r_ order influence of the expanded isentropiclly to free stream conditions and unsteady density ratio can be neglected (i.e. ignoring is in general less than the centerline jet velocity at compressibility effects and cross correlation terms).
the exit of the nozzle as will be shown later in the Figure 8 illustrates a pulsed jet having a 20% duty cycle text. The momentum coefficient can be expanded and the potential demand for actuator authority, (the to become: magnitude required to influence for flow field). Even if an ideal pulsed jet can be created, it wilt become distorted at the exit plane due to the compressible
(Eq.4)
effects that are related to the volume of the plenum. As Since the jet velocities are expected to be this distortion occurs the demand on time accurate compressible, the density ratio will be based on the measurements at the slot exit become more important in Mach number related to equation 3. Figure 7 understanding the physics of the pulsed circulation system.
AIAA 1= Flow Control Conference 4 AIAA 2002-3157 1.2 Perhaps a more appropriate parameter is a kinetic 20% DUTY CYCLE energy based correction that is expressed as: /.. \ (Eq.14)
CD( ou j
o8 RMS: 0.447 / To include an additional penalty for mass intake (ram Cp. 0.6 _ effect) results in: 0,4 +C,, U---_J/+ AVERAGE: 0.2 CD(EQUIV) = CD(BAL)(ZU= / C"(-_'-j / (Eq'15) 0.2 .......................................................
Comparisons of these correction techniques are shown 0 , _____.,--- .0.2 0.0 0.2 0.4 0.6 0.8 1.0 1.21.4 in figure 9 by highlighting the components being added TIME to the measured drag. Nominally the momentum Figure 8. Example of a theoretical pulsed jet coefficient is small for cruise conditions resulting in a minimal impact of the correction technique. As the Experimentally many researchers are not momentum coefficient increases for high lift conditions always able to make detailed and accurate at low speed, the selection of the correction technique measurements at the jet exit due to large velocity becomes more critical. Of the three equivalent drag perturbations and/or small scales. Measurements techniques, the kinetic-energy based term coupled with in the plenum or upstream of the model are often the mass intake penalty is the most conservative and used to quantify mass flow and jet velocity to best represents the physics related to the drag and will "standard" or free stream conditions. This also be used throughout remainder of this text unless enables the researcher to manage the density ratio otherwise noted.
more effectively. If the mass flow measurements are made far enough upstream, the perturbations are damped and the total mass flow conditions are captured. Equation 1 then becomes: CIa Uj +CI_ U= _ /
0. (/ t/
2U® Uj
(Eq.,O)
o,- q(b)(C)
where the mean and fluctuating velocities can be obtained from equation 11 and 12.
0.2 U j= 2¥:I(Tr'ucT)1 ( P= / T- 0 0.05 0.1 0.15 0.2
y-1 -t_) j (Eq.11)
ClA Figure 9 Comparison of equivalent drag correction I 7-1 techniques CFD Analysis y - 1 _PLOCAL ) The NASA LaRC Full Unstructured Navier- In addition to eliminating the issues associated with Stokes 2D code (FUN2D) was used for the CFD density ratios, this approach also eliminates the computations. The flow solver is a node based, errors associated with the measurement of the slot implicit, upwind flow solver used for computing flows height.
around single or multi-element airfoil configurations with The influence of the Coanda jet on the non- unstructured grid. le The governing equations are the dimensional lift, drag and pitching moment are time-dependent Reynolds-Averaged Navier-Stokes realized in the forces created on the airfoil. It is (RANS) equations in conservation-law form, which are necessary to define an equivalent drag for an airfoil integrated in time to obtain a steady state solution. The that has a potential for creating thrust. This enables Spalart-AIImaras turbulence model 19 was used in this blown sections to be comparable to conventional investigation and all computations assume fully airfoils and avoids lift to drag ratio (airfoil efficiency) turbulent flow. Boundary conditions that enable blowing from going to infinity as the measured drag capability were integrated into the code. Steady jet approaches zero due to blowing. 17 A single force- calculations were performed over a range of slot based coefficient simply includes the momentum heights, blowing rates, and angles of attack.
coefficient to the measured drag coefficient.
The grids were generated with advancing front type point placement with iterative local re-meshing for CD(EQUIV) = CD(BAL) +Cp (Eq13) grid quality improvement. 2° 21 The outer boundaries AIAA 1= Flow Control Conference 5 AIAA 2002-3157
were treated as characteristic inflow and outflow
The GACC airfoil used throughout this study was a surfaces and extended 20 chord lengths in all modified GAW(1) airfoil. This modification occurs only directions from the leading edge of the model. No- on the under surface near the trailing edge to slip viscous boundary conditions were prescribed on accommodate the 2% r/C Coanda surface. The chord the airfoil surface and the blowing boundary reference line used for angle of attack extends along the conditions were applied to a vertical plane internal original GAW(1) leading and trailing edge references.
to the model.
For the GACC airfoil this 2,0 For a chord Reynolds number of 533000, FUN(PD) the turbulent minimum normal wall spacing on the airfoil was set to 2.0 x 10 .6 based on a chord of 1.0.
G_ WSU The grid generated for the GACC airfoil consisted of 90,582 nodes with 1540 nodes on the airfoil surface 'L'ockheld CL iF'r 1.0 and the duct walls. Figure 10a shows the _e 41_ , (l_j' unstructured grid generated around the GACC model. Figure 10b shows an enlarged view of the grid in the duct region, highlighting the blowing boundary condition that coincides with the exit plane diffuser.
O_ AOA Figure 11 .FUN(2D) validation using GAW(1) experiments corresponds to the same leading reference and a trailing edge reference that is tangent to the upper Coanda surface coinciding with the upper surface jet exit plane.
Initial studies of the Coanda turning revealed the need to model the internal plenum due to discontinuities at the slot exit and difficulties with convergence. The internal GACC plenum to jet exit contraction ratio varied from 10:1 to 20:1 depending on the slot height. The magnitude of the jet velocity was a) Unstructured airfoil controlled at the internal boundary corresponding to the exit of the actuator.
The predicted lift performance of the GACC airfoil shown in figure 12 is consistent with similar I supercritical airfoils. 2' Sequences of flow fields with varying blowing rates are shown in the appendix. The visible boundary shown in this sequence is I FLOW CONTROL [ BOUNARY S.0 I t--cp:0.1s= GACC 2-D AIRFOIL 4.0 r--Cl=: 0.091 I , $.0 : BLOWING CLIFI- Ct=: 0.031 b) Enlarged view of GACC trailing unstructured grid 2.0 +.Ctm: 0.021 NO Figure 10 FUN(2D) grid for GACC airfoil high lift _.._r_e BLO WInG .CIA: 0.013 ..4P--" configuration with a slot height of 0.020" 1.0 The initial expectations of the CFD effort 0.0 : : t I were built around using the code as a toot for 0 $ 10 15; ZO designing the experiment. To gain confidence in AOA the application and use of the code in the GACC Figure 12 GACC predicted performance using FUN(2D) research program, a test case was run for a GAW(1) airfoil. A comparison of GAW(1) lift representative of the wind tunnel walls and shows the coefficient data with two independent experimental potential wall interference effects. As the jet velocity studies" _ showed favorable results, (figure 11).
increases, the Coanda turning angle also increases until AIAA 1= Flow Control Conference 6 AIAA 2002-3157 wall-bounded jet decelerating as the low momentum it reaches a limit imposed by the GACC's lower slot outer flow is entrained near the jet exit. It then briefly exit. The Coanda jet penetrates the low momentum accelerates and remains attached to the Coanda wake and deflects the streamlines near the trailing surface until confronted with the forward facing step of edge. As the jet velocity continues to increase, the the lower jet exit plane. The large suction peak at the Coanda jet penetrates the flow at a fixed angle. The vectored jet departs the Coanda surface into the trai!_q edge significantly oncoming flow field, creating a negative thrust (e.g.
_ C_._ 0.162 similar to a thrust reverser). The degree of penetration is dependent on the magnitude of the jet exit velocity and free stream velocity. As the jet
".0lY \
penetrates the flow field and turns the streamlines the effective pneumatic flap is created. Cp -lO.O _ _ \ I _""J" _ '_ SEPARATION The corresponding pressure profiles shown
._.o_ J. _/ \ _,\ (,-,67 °)
in figure 13 highlights the pitching moment created -25.0 o, ....
AOA 0.0 -20.0 ] ,,s.O - h 0.010" _ ....
0 :310 60 90 120 150 180 _ u./o-" COANDA TURNING ANGLE (_) -15,0 Figure 14 Comparison of two pressure profiles along the Coanda surface Cp -lo.o contributes to increasing the lift, drag, and pitching -8.0 moment.
0.0 The Mach number profiles at the jet exit are consistent with a developing internal nozzle flow field as 5.0 shown in figure 15. If one uses the static pressure at 1.0 0.0 0,2 0.4 0,6 O.B X/C the exit plane on Coanda surface (peak pressure shown Figure la GACC pressure profile for varying in figure 13) the resulting Mach number will be momentum coefficients (0<Cp<0.2), Mach 0.1 inconsistent with the centerline Mach number of the Coanda jet. The use of the jet velocity based on the by the large velocity at the jet exit. Looking closer at the trailing edge reveals a large pressure difference across the jet exit shown in figure 13.
This is characteristic of a wall-bounded jet. Caution must be given in using the pressure data from the Coanda surface to estimate the jet velocity at the exit as will be discussed.
_ 2 5 D 0 k UPPER JET EXIT\
== o__o=i _
'-° SORFACE \ f ......
TRAILING EDGE --_-'--'J _ '.,.,U_LJ_ j
"_ l SURFACE / 0._I_1 "
LOWER .j 0.00 0.,50 1.00 Cp -10XI, SURFACE j/ MACH Figure 15. Jet exit Mach number profiles for GACC h:0.010", free stream Mach:010.
.OWER / f
-=T EX_ISE_T,O./
0.0" _ : 1/ : i ! I I I centerline Mach number to calculate the momentum coefficient will result in an over-prediction of the momentum due to losses associated with the growing 0.970 0.975 0.980 0.965 0.990 0.995 1.000 X/C internal boundary layers. This is particularly true for smaller slots where the nozzle boundary layers are a Figure 13. Trailing edge pressure distribution along the Coanda surface for C_ 0.162, Mach 0.1, AOA significant portion of the jet flow field. Therefore, integration of the jet profile is necessary. Nominally this 0.0, and h/C 0.001.
will result in a lower jet velocity magnitude. This may explain why using free stream static pressure (isentropic Examining the same profile relative to the Coanda surface shown in figure 14, reveals the AIAA 1stFlow Control Conference 7 AIAA 2002-3157 expansion of jet) to calculate the Coanda jet velocity system that is internal to the actuator showed a 15% has been historically successful in momentum reduction as the frequency increased from 25Hz to calculations.
100Hz figure 17.
7,5- I Exi_erlmental Setuo The development of the GACC test program is based on a pulsed circulation control concept.
Actuator authority and frequency response are characterized by state of the art high-speed valves ,":o and are the cornerstone to the model development.
The model requirements were established to provide baseline and unsteady circulation control data for proof of concept and code validation of pneumatic flap and control surface concepts.
0.0 Duty Cy©lo Actuator Description Figure 17 Actuator mass flow authority of a high-speed Studies that have attempted pulsed solenoid valve operated at different duty cycles and circulation control have historically been limited to frequencies. Inlet pressure 200 psig.
rotary or shuttle valves and have had limited authority and frequency response. Requirements Assuming that a CI_ of 0.1 is required to achieve the for the current study were to extend the frequency desired lift, 100 SCFM wilt be required of the actuator survey to 200 Hz with peak velocities that approach system, see figure 7. Figure 18 shows the electronic sonic conditions. Pulsed actuation can be timing associated with a typical injector. The hot wire generated through valving the flow path to the location had to be moved one inch downstream as the trailing edge. While it is convenient to place the high velocity pulsed stream kept breaking wires.
valve system outside the model, the frequency S response will decay as a result of the volume and HOT WIRE flow path leading to the jet exit at the trailing edge.
_-- 825 Ws To minimize this decay and avoid unwanted 3-D effects, it is important to locate the valve system as close to the jet exit as possible. This requirement lead to the development of a high-speed solenoid ACTUATOR o,"E N SIGNAL operated valve system that would be distributed DRIVER HOLD GENERATOR " along the span of the GACC model.
"" CLOSE Figure 16 shows an example of a high- 0 0,2 0.4 0.6 0.8 speed solenoid operated valve, (i.e modified fuel Time (Im¢) injector). This valve utilizes a piston that seats into Figure 18 Comparison of the timing of the pulsed an 0.010-inch orifice, and is cycled from a full open actuator system, frequency: 20 Hz, DC 0.50, Inlet to a full closed condition. This enables the valve to pressure 200 psia.
generate a pulsed flow with variable frequency, duty cycle, and velocity magnitude.
Evaluating the general performance of a single actuator (tables 1 through 3) one can see that 20 actuators will SOLENOID HOUSING be required to operate simultaneously at 5 SCFM to ORIFACE meet the 100 SCFM imposed by the lift requirement. It is clear that not all frequencies and duty cycles meet m..m 0 this requirement.
!
/ 'O' RING SPRING PISTON Actuator Rapid Diffuser Description The flow field out of the actuator is a small Figure 16 Breakdown of a high-speed solenoid valve.
diameter circular high-speed jet. To avoid unwanted 3-D span-wise flows it was necessary to design a rapid The calibration of a single actuator provided diffuser that could be integrated into the GACC model.
information that was used to size the high pressure The objective of the diffuser is to transition from a air delivery system and quantified the number of circular, time-dependent, high-speed jet to a low speed actuators required to meet the mass flow 2-D, uniform jet. This was accomplished with a 10:1 requirements necessary to achieve the target lift area ratio diffuser shown in figure 19.
coefficient, The mechanics of the piston-spring AIAA 1 = Flow Control Conference 8 AIAA 2002-3157 components respectively. These frequencies were avoided by driving the pulsed actuator system at frequencies above or below the predicted natural frequencies. The axial component was easily excited, even with steady blowing conditions. This limited the test matrix and increased the errors associated with drag measurements for certain conditions. Figure 21 shows the assembled GACC model system.
..... .
20 ACTUATORS GACC W/ w/UPPER RAPID DIFFUSERS SKINS REMOVED Figure 19. Rapid diffuser with internal guide vanes for GACC Actuator.
Model Description INSTRUMENTED BART The GACC model was designed around the GLASS SIDEWALL COANDA wl PRESSRE PORTS pulsed actuation system and sized to fit into the TRAILING EDGE LaRC BART wind tunnel. The model is a 17% thick supercritical airfoil that has a 9.4-inch chord and a 28-inch span (aspect ratio 2.98). There are 20 AOA independently pulsed actuators distributed along the DRIVE span of the model, figure 20. The model is mounted vertically through the tunnel floor and spans the 5 COMPONENT entire height of the tunnel.
BALANCE 20 ACTUATORS TRAPEZE w! DIFFUSERS Figure 21. Photograph of GACC model assembly An electronically scanned pressure system was used to measure 52 steady model pressures and 150 wind tunnel wall pressures. Two 32-port modules were mounted internal to the model to avoid balance INSTRUMENTED interference. The trailing edge Coanda surface was TRAILING EDGE instrumented with 40 thin films and 13 unsteady surface COAN DA pressure transducers shown in figure 22.
SURFACE COANDA Figure 20. GACC Model w/upper surface removed SURFACE to expose internal actuator system TO accommodate the 5 component balance system the model was cantilevered and a 0.125- inch gap was left at the ceiling to avoid fouling. The
/
5-component GACC strain gage balance was INSTRUMENTATIC PRESSURE designed around the lift and drag performance SPAR WAFERS predicted by the CFD analysis. Since the model was Figure 22. Trailing edge instrumentation mounted vertically, the gravity direction was package for all test conditions.
ignored. A description of the balance limits is shown in table 4.
Each pressure wafer had multiple independent The model design also accommodated one 0.020" 5-psig pressure sensors bonded to the surface.
high pressure and two low-pressure air supply lines During the course of testing it was discovered that a that were coupled to the air delivery system through small pressure leak developed at the interface of a a trapeze _4. The trapeze was fabricated with pressure wafer and the instrumentation spar. This flexible hoses that minimized the forces transmitted biased the mean pressure measurements but was to the model. The tare forces associated with the assumed to have a minimal impact on the fluctuating air supply system were statically calibrated then pressure measurements due to the slow leak rate.
applied real time to the force data.
The GACC air supply system design, figure 23, The balance had natural frequencies of 12.7 was based on a C_ that would be limited to 0.2 or less.
Hz and 78.6 Hz for the axial and normal AIAA 1= Flow Control Conference 9 AIAA 2002-3157 This corresponds to a maximum mass flow of 150 ExDerlmental AnalyslF SCFM (see figure 7) and enables the researcher to The evaluation of the GACC performance is test through predicted lift coefficients of 4.5 for broken into three separate efforts, 1) baseline high lift different slot geometries. Two of the air supply performance using steady blowing, 2) cruise systems were designed to provide independent performance using simultaneous upper and lower control of both upper and lower steady blowing.
steady blowing, and 3) high lift performance using Each of these independent systems was limited to pulsed blowing. Two GACC slot geometries were 30 psig of controlled steady air to the isolated upper tested (h/r: 0.0533 and 0.1067) and selected data will be and lower plenums. The third air supply system presented. Wall interference corrections have not been supplied the actuator manifold with a regulated applied to any of the experimental data.
pressure that ranged from 50 psig to 200 psig. The mass flow for each of the air supply lines was 1) Baseline High Lift Performance independently measured at the exit of the air supply To capture the details of the physics associated system with turbine type flow meters. The actuator with circulation controlled concepts it is often beneficial line was buffered with a surge tank to avoid to step back and look at the global picture. A flow pressure pulses back to the flow metering system. A visualization experiment was performed to provide real detailed description of the GACC model design can time flow analysis. A separate and smaller GACC be found in Reference 13.
model was designed and built for water tunnel applications. Figure 24 is an example of a hydrogen ..................... CONTROLLER bubble flow visualization technique used to evaluate the flow tuming of the GACC model. Movies highlighted =,._ "i_ _. "_f_ | BOOSTER both positive and negative high lift characteristics as well as the cruise condition.
_/"l -- _ -_ _I_ _°.,.°L_.
I UH_I_E" / ==oL_rm_ VOLUME Y'U_IBINI= FLOW METER RE-_O,k._,.AT_ BOOSTER _t OW M_C'IEH Figure 23. Schematic of GACC air supply system No Facility Description (BART) The NASA-Langley Basic Aerodynamics Research Tunnel (BART) is a flow-diagnostic facility that specializes in the acquisition of detailed data for the development and validation of CFD models and advanced flow diagnostic techniques. Its flexibility and advanced measurement capabilities are often utilized to investigate the fundamental characteristics of complex flow fields about various vehicle configurations.
(b) Upper Surface Blowing BART is a subsonic, open-return wind Figure 24. Hydrogen bubble flow visualization of the tunnel with a closed test section 28 inches high, 42 GACC airfoil.
inches wide and 120 inches long. During operation, air is drawn into the tunnel inlet under atmospheric The flow visualization experiment highlighted several conditions by a 9 blade-11 stator fan. The fan is important features of the GACC model, such as powered by a 125 horsepower, alternating current • 3-D slot flow can be generated without appropriate motor coupled to a magnetic clutch. Air passes internal flow conditioning resulting in inefficient flow through a honeycomb, four anti-turbulence screens, turning, and an 11:1 contraction before entering the test ° velocity ratios near 1 are the most efficient for the section. The maximum velocity at the test section dual blowing cruise conditions, and entrance is 186 ft/s, which corresponds to a unit ° tunnel walls can become separated with high Reynolds Number (Re/if) of 1.13 million and a degrees of flow turning that are characteristic of dynamic pressure of 40 If/ft _. The turbulence circulation control wings.
intensity varies from 0.03% at 50 ft/s to 0.09% at tunnel maximum velocity.
AIAA 1 = Flow Control Conference 10 AIAA 2002-3157 Due to the lack of confidence in the unknown BART High Lift Performance trailing edge pressure and the associated error The high lift performance will be discussed magnitudes, pressure data could not be used to obtain in two separate sections; a) positive high lift (Upper force data that would compliment the balance surface blowing) and b) negative high lift (Lower T surface blowing) AOA 0.0 L -10.0 -_ _ EXP a) UDDer Surface Blowine ""_,/f_ _ (BART) The baseline performance of the GACC :: _ _ / CFD model is characterized by the lift and lift efficiency (L/D) shown in figures 25. The GACC experimental Cp -5.o • "_ ;FUN2D) lift results evaluated at an angle of attack of zero will provide lift augmentation, ACL/ACI_=50. This is 0.0 , l _ , , .2,----_ consistent with other small trailing edge CCW airfoil experiments.
5.0 0.0 30.0 00.0 0(I.0 120.0 150.0 180.0 4.0 COANDA TURNING ANGLE (t) Figure 25 Comparison of CFD and experimental Coanda pressure distribution for the GACC airfoil, h 3.0 t 0.010", nominal CI_ 0.06 G_ 2.0 data. This also made corrections to the momentum CLIFf coefficient (equation 5) impractical since the measured pressure at the jet exit was unreliable. Therefore, the momentum data for the CFD and experimental results are both based on equations 1 and 3. This would 0.0 0_" _ . I . , suggest that there will be a potential difference in the -10 -5 0 5 10 15 described lift performance for the different slot heights.
AOA (Degrees) CFD predicted the GACC experimental Figure 25 Lift characteristics of the 2-D GACC airfoil performance trends for the small slot configuration very well, as shown in figure 26. As noted above, the lift There is a favorable comparison of the CFD performance for the different slot heights should be and experimental pressure data as shown in an different as demonstrated with the CFD data shown in example in figure 26.
figure 26. However, this trend is not realized in the experimental data and is not understood at this time.
-15.0.
4.0 ..........................................
AOA 0.0 AOA: 0.0 ° _" J" "#""" .110.0 EXP CFD (BART) (FUN2D) 3.0 _._¢_/'#'_'_'" - " ....
SYMBOLS SOLI D LINE
_c,. ,,, ,e.a: "x
ACp EXP "f_2 'Jr" SLOT HEIGHT Cp -S.0 _t_:
c.. =.0 \ _, (,_,,)
m_"m _ EXP 0.0101 0.0 _ . . • I . • , .... I . . .
I.o If" . ceo o.01oj 5.0 J- O_ • : . t 1.0 0.0 0.2 0.4 0.6 0.8 X/C 0.000 0.025 Oc_D _ 0._/$ 0.100 Figure 26 Comparison of CFD and experimental Figure 26 Comparison of CFD and experimental Lift pressure for the GACC airfoil, h 0.010" CI_ 0.06 characteristics of the 2-D GACC high lift airfoil configuration.
The differences in the trailing edge pressures are highlighted in figure 27. While the trends are The drag polar for the GACC high lift similar, the confidence in the experimental configuration at zero degrees angle of attack is shown in pressures is limited due to a slow pressure leak into figure 27. The general trend is consistent with the reference plenum that developed during the performance characteristics of traditional high lift test.
systems that vary in angle of attack. However, the CFD results are lower than the experimental data and may be AIAA 1= Flow Control Conference 11 AIAA 2002-3157 a result of the experimental wall interference, at a momentum coefficient of approximately 0.05 with a balance errors, mass flow measurement errors, maximum negative lift of -1.0.
and/or errors in setting the slot height of GACC 1,0 .
model.
AOA = 0.0 4.0 0.5 _ Cov:_/= 00098 o - "`41+ _'_ L_, .G AOA: 0.0 _-:+-:" _ t, II L 0.0 . . . II. . _ ; ,, II' 3.0 Ct.iFr o -0.5 ,_I _ _._ + "_' SLOT HEIGHT GuN 2.0.
-1.0 LEADING EDGE 1.0 STALL ,_ I ÷o_'o+ [.CFO Ulo -1.5 0.O25 0.05 0.075 o._ C_ 0.0 .... ; .... : .... : .... : .... , Figure 27 Lift performance of GACC 2-D airfoil negative 0.00 0.10 0,20 0.30 0AO 0.50 ComAs lift configuration (Lower Surface Blowing) (AOA:0.0) Figure 27. Drag Polar for GAG@ 2-D high lift airfoil configuration using Equivalent Drag (0<CtLt_-'_). 10) The drag polar of the negative lift configuration decreases from the un-blown baseline, figure 28.
Comparing the minimum drag of the upper blown high The airfoil efficiency for the GACC airfoil at lift configuration shows a small improvement in the zero degrees angle of attack is shown in figure 28.
minimum drag for the negative high lift configuration.
The trends can be compared to conventional high Using lower surface blowing one can optimize the cruise lift systems described by C.P. van Dam. m While the drag for a given lift coefficient, (eg. CL 0.25 and overall airfoil efficiency of a conventional airfoil Co.i, 0.0098 are typical for a general aviation aircraft) system is higher than the GACC airfoil shown here, there remain questions regarding the uncertainties %0 .............................. .................................................
related to the kinetic energy components added to the balance data.
0.5 •. _, AOA: 0.0 _ SLOT .EIGHT 0+:i,: .... : .... : .... : .... :.... :....
, . _ _k <Inches) CLI t'l z '_ i,_ExPo-tool IJJ M "- , II CFO 0.020 -I.0 INCREASING
-
-1.5 ......
0 0.05 0.1 0,15 0.2 0.25 0.3 0.35 CD_O 10, INCREASING _Z_'" Figure 28 Drag Polar for a GACC 2-D airfoil negative lift configuration (Lower Surface Blowing) (AOA:0.0) 0 .... : .... : .... : .... : .... ,' 0 1 2 3 4 5 Cti_ C) Dual Blowing Figure 26 Airfoil efficiency (I._) of a 2-D GACC high The combined or dual blowing configuration of lift airfoil configuration the GACC airfoil is targeted at understanding the minimum blowing requirements for cruise conditions of a b) Lower Surface Blowing CCW airfoil. For cruise conditions the ratio of the jet A unique feature of the GACC airfoil is its velocity to free stream velocity is a critical parameter.
ability to generate negative lift by blowing from the As the free stream velocity increases to high subsonic lower slot. While one may question usefulness of speeds, the jet velocity required for optimal cruise will this configuration, it has the potential to make rapid also increase. It is important to keep the interactions flight path corrections depending on the frequency with any shock formations to a minimum. For this series response of the system. The momentum of experiments the free stream velocity was kept low to requirements of the negative lift configuration are avoid compressible conditions. Figure 30 highlights the shown in figure 27. The stall characteristics occur uncorrected drag and the measured thrust for two slot configurations at zero angle of attack. The larger slot AIAA 1 " Flow Control Conference 12 AIAA 2002-3157 configurations. This is due to the high drag values develops more thrust vectoring as a result of increased mass. being influenced by the kinetic energy and momentum effects.
1.60 r S1_01 40.0 I 1.40 h (eNCHES) I ¢: 0.01 I • Ulet/uo o 1.20 • 0.0_ 30.0 -- eo.oo *O,M m•.
1.00 ,-% ,O,IWI !
|1.11 _2.3E* _-_,. 20.0 -- CLIF T 0.80 ', i 0.60
, I
IO.O
Y
0.20 .... ] 0.40 _i t W7 0.00 t Im 0.000 0.020 0.040 0.080 0.060 0.O40 0.020 I -10.0 CDEAG (BAt } -10_0 -5.0 0.0 5.0 10.0 AOA (Degrees) Figure 33. Dual Blowing Lift to Drag ratio (h: 0.01) Figure 30 Dual Blowing Influence of two slot heights (Balance data) The drag polar shown in figure 31 illustrates a d) Pulsed Blowina significant benefit of the dual blowing configuration.
A 38% reduction in drag occurs at a -4 ° angle of The effectiveness of pulsed blowing on the attack and velocity ratio of 1.2 shown in figure 32.
performance of the GACC airfoil is dependent on the 1.5 efficiency of the actuator system. This system must include the actuator performance, diffuser performance, and the response of the internal volume prior to the jet exit as well as the external time dependent Coanda effectiveness. Ideally the time dependent Coanda
i .o 1 u._,fuo=l 2 j _gl_ I l
response would resemble the steady state blowing series shown in the appendix. This would assume a
.i
perfect square wave response at the jet exit. The reality • ,.., t ' . . .o.om2L , 0.0 of a perfect square wave diminishes with the complexities of the actuator system.
Uj C U.
The response of the state of the art high-speed 2 L -O.5 0.10 0.00 0.02 0.04 0._ (}.(18 actuator valves used for this study does not generate a CWULG perfect pulsed _ waveform as typified in figure 16.
Figure 31. Dual Blowing Drag Polar (h: 0.01) Transmitting the pulse through the nozzle and into the nozzle exit distorts the waveform further as shown with 0.10 / thin film data located at the nozzle exit. For the low ute_Jo A frequency pulsed jet, the effect of duty cycle is shown in 0.08 -- ,,.o.u _ ,c, • 1 34. The peak amplitude for the low duty cycle conditions (20% and 30%) does not reach the maximum
!'17
0.06 -- output performance of the actuator system. This result CD is caused by the actuator valve being closed before the 0.114 plenum and actuator volumes have had time to be fully pressurized. Once the valve is given a close command the plenum remains pressurized and continues to bleed 0,02 U_ U.
air through the jet exit until the plenum pressure reaches ambient conditions.
0._ -6.00 0.00 5.00 10.00 15,00 -I0,_ As the drive frequency is increased, figure 35, AOA (Degrees) the rise time or valve opening distortions increase. For Figure 32. Drag performance of the GACC airfoil for the closed portion of the duty cycle, air continues to different velocity ratios.
bleed from the plenum until the open command is given, resulting in the jet velocity not going to zero. This The trends in dual blowing efficiency, shown in process limits the mass flow to the jet exit as indicated figure 33, are low compared with the peak by an overall reduction in the peak velocity. In spite of efficiencies of conventional airfoil cruise AIAA 1_ Flow Control Conference 13 AIAA 2002-3157 1.2 ¸ the limitations of the actuator system, the peak velocities do approach sonic conditions.
1.Z
, 'l ...... li I
vAmwm.E OUTYCY_E --20% _ _- '50_ ---60_ --70% --R0%] o.
°.6 : - .. '
o.e i E-Etd_v 0.6 0 0.02 0.04 0.06 0.08 0.1 TIME (Sec) Figure 36 Time trace of pulsed let along the Coanda 0 0.02 0.04 0.06 0.08 0.1 surface (Driver frequency 35 Hz, 50% duty cycle, and TIME(See)
h 0.020")
Figure 34 Normalized thin film time history for pulsed CCW at the slot exit. (h: 0.020 and Driver CLI Frequency: 35 Hz) 0.01 INCREASING DECREASING DUTY CYCLE EFFICIENCY ;i _=4.5 1 (40% < OS < 80%) -....
0.001 WIND OFF O,8
/ i' _ti
=t -0.0 0.0001 , O.6 0.000O1 10 100 1000 10000 0.4 FRE(]{JENCY (Hz) Figure 37 Sp_ra of the pulsed jet in the _cini_ of the _ti_ I+ 0.2 slot exit, (35 Hz, 50% duty cycle, and h 0.020"1 O.O2 0 0,04 0._ 0.08 0.1 Comparing the pulsed and steady lift TIME (sec) performance of the GACC airfoil, figure 38, a distinct Figure 35 Normalized Thin Film Time histories for improvement in can be seen. For a given lift coefficient pulsed CCW at the slot exit, (h" 0.020 and Driver of 1.0, a 48% reduction in mass flow is realized for a Frequency: 100 Hz) 20% duty cycle. As the duty cycle is increased the performance benefit decreases. Comparing the lift Figure 36 illustrates the authority of the performance of the pulsed and steady CCW at a fixed pulsed jet on the Coanda surface. The magnitude mass flow of 25 SCFM results in a 35% lift of the jet decreases as the flow exits the nozzle and improvement.
moves along the Coanda surface. It is important to recognize the structure of the pulsed waveform is 2.5. : maintained until the flow separates from the Coanda surface.
As the pulsed flow exits the nozzle, a large increase in the turbulence characteristics are observed. The spectra of two surface thin films Cu_ near the nozzle exit are shown in figure 37. These 2.0.1.5. _ __"_10% "_.
1`0, data are consistent with conditions measured for both slot configurations. For low frequency pulsed jets, the frequency characteristics of the jet can be separated into pulsed and turbulence regimes. The 0`0 transition from the pulsed to the turbulence regime seemed to be independent on the jet velocity and mdot (SCFM) occurred near 300 hz.
Figure 38 Comparison of pulsed and steady circulation control, (Frequency 35 Hz and varying Duty Cycle).
AIAA 1= Flow Control Conference 14 AIAA 2002-3157 the actuator itself. Continued research is necessary to quantify the overall system time dependent response of Concludln_a Remarks the GACC airfoil including the leading edge and internal plenums. Modifications to the internal flow path to The steady and pulsed aerodynamic improve the pulsed system authority will be necessary performance of the GACC airfoil has been to increase the overall lift performance.
demonstrated using both CFD and experimental The GACC airfoil has been proven to be an methods. The FUN2D code used in this study excellent test bed for the multi-functional circulation predicted the performance trends of the experiment control study that can operate as a high lift system, a very well and proved to be an excellent tool for pneumatic aileron, and a high-speed air brake. Follow- identifying flow features that were subsequently on testing is expected to improve the database for CFD probed in the experiment. The CFD results were validation and the understanding of the flow physics also a valuable asset in interpreting the related to circulation control concepts.
performance characteristics related to the Coanda jet and the GACC high lift configuration.
INPUT PRESSURE: 100 PSIA The baseline performance for the GACC FREQ DUTY CYCLE high lift configuration is similar to other circulation control wings described in the literature. The 2(P/= 25% 50% 75% 80% performance of the pneumatic aileron in cruise 25HZ 0.8 1.0 2.1 3.1 3.3 conditions (i.e. dual blowing) showed significant 50 HZ 0.7 1,0 2.0 3.1 3.3 improvements in drag over the unblown baseline 100 H2 0.7 0.9 2.0 3.0 3,2 configuration. Application of the GACC airfoil to a 150 H2 0.5 0.7 1.8 3.0 3,3 specific general aviation or personal vehicle may or 200 H2 0.0 0.2 2.0 3,1 3.7 may not be practical due to the large baseline drag.
STANDARD CUBIC FEET PER MINUTE It is believed that the airfoil can be optimized to actuator performance at inlet pressure of Table 1. Single reduce the baseline drag but is not within the scope 100 psi(.i of this flow physics study.
This study has provided extensive details in INPUT PRESSURE: 200 PSIA the region of the Coanda surface and has identified FREQ DUTY CYCLE issues related to the modeling of the Coanda jet.
20% 25% 50% 75% 80% The isentropic expansion of the conditions at the jet 25 HZ 1.5 2.0 4.2 6.4 6.8 exit quantifies jet velocities that are typically lower 50 HZ 1.3 1.7 3.9 6.1 6.6 than the integrated velocity at the jet exit. The 100 HZ 0.7 1.3 3.5 5.7 6.1 magnitude of this error is a function of the slot 150HZ 0.0 0.1 2.7 5.0 5.5 height and the internal development of the jet exit _)00HZ 0.0 0.0 2.7 5.1 6.1 profile.
;TANDARD CUBIC FEET PER MINUTE The difficulties in experimentally measuring actuator performance at an inlet the jet exit profile for the GACC airfoil is realized in Table 2. Single the spatial resolution of the trailing edge. The pressure of 200 psi{] INPUT PRESSURE: 300 PSLA physics of the jet velocity at the trailing edge, FREQ DUTY CYCLE particularly at the jet exit are critically important in 20% 25% 5_ 75% 8_ understanding the performance benefits of steady 25HZ 2.1 2.7 6.2 N_ N_ and pulsed pneumatic control of the GACC airfoil.
50HZ 1.4 2,0 5.5 9.1 N_ The measured Coanda surface pressure profile I_ HZ 0.0 0.8 4.5 7.9 8.4 identified a pressure peak at the slot exit that was 150 H2 0.0 0.0 2.7 5.1 5.8 2_ H2 0.0 0.0 0.8 5.0 6.1 consistent with CFD results. Using this pressure to STANDARD CUBIC FEET PER MINUTE quantify the jet velocity and momentum coefficient did not realize a collapse of the airfoil performance Table 3. Single actuator performance at an inlet data. Further study is necessary, particularly for the pressure of 300 psig compressible jets created when using small slot Load heights.
Component The performance of the pulsed blowing Normal 100 Ib system realized a 50% reduction in required mass Axial 5 Ib flow for a given lift coefficient. Variations in the duty Pitch 400 in-lb cycle at a given frequency highlighted the Roll 1200 in-lb controllability of the performance with small bursts Yaw 40 in-tb of high-speed air. The study of the pulsed air Table 4. Load limits of the GACC 5-component strain delivery system identified losses related to the gage balance plenum volume and the performance requirement of AIAA 1 = Flow Control Conference 15 AIAA 2002-3157 A DDendlx Mech uech 0+OO0o 0+029o 0+|200 I+0_0 o,sooo O+OOOO *" 0.04O0 0+26OO OJIOOO 0+IOOO O+;'OOO 0,?OOO O.72OO O.72OO
OiooO Ujet: 0.00 0,61o0
o+2600 O 1 _ 0 0 0,40O0 OJOOO
Curr: 0.66
O.sooO 0_Slo0 03|OO 0,S2OO O.qoo 0,mIOO 0.441)0 O.44OO 0.4000 0,4O0O 0_IO0 0.24OO 0+_IOO 0.$OOO 0.2100 0.2100 O+24OO 0,24oo ++ O,2OOO _=+ O+2O00 0+OOOO i Q++OOO i 0.11OO 0,26_ 0._ _NOO
Ujet: 308 ft/s
C,,,: 1.39
i
Ms_ O.8OOO 0.OO00 0.7OOO 0.70OO O.72OO 0.7_00 O.IOOO 0.OO00 0.26OO 0.1IIOO O.OOeO
Ujet: 379 ft/s
0.OO_ O+SllO0 0.OOOO 0.520O 0.S2OO
1.83 p+411OO
0.4OO0 O+I4OO 0.4400 O+4OO0 O.40OO O+31OO O,=OOO 0.32OO 0+2loo O.2400 OJ_O 0.121_ o+lml_ 0,1_ 0.OOOO O+04OO O,O4OO D.O0OO 0.OOnn laeJc_ 0.gooO o.es9o 0.1OOO 1,0000 o.oooo 0.411OO _ 0.84OO + o+26oo O.OO0O o.looo 0.7800 o.7ooo 0.7_0 o.72oo o.18oo OJMIOO 0.1404) oJ4oo 0,IGO0
o.oooo Ujet: 460 ft/s
0+SOO'0 4)+1ooo O.S2OO o.iNe
CuFf: 2.30
o.41oo 0.41100 0+44_ oA4oo O,4OOO O.ILI_ O.311OO 0.3100 0.+3200 0.32OO 0.2100 0.2100 O.,1400 0.26OO 0.29OO 0,161_ 0.1 2OO 0,G2OO 0,OOOO i 0.20OO 0.1600 O,O4OO O,O4OO 0,_GD 0.OOOO Appendix 1 Influence of Coanda Jet on the Mach number distdbution AIAA 1" Flow Control Conference 16 AIAA 2002-3157 -I1
g_
I- B ¸ _I m_! 11 iiiiiiii_!i_i_iiiiiiiiiiii_ iiiiiiii_,ii_ii!iiiiiiiiiii_.
iiilliiiiiiiiiiiiiiiiiiiii!
ii!! ¸ I"O ,,.q ,5 Rose, R.E., Hammer, J.M., & Kizilos, A.P., "Feasibility I:_eren ¢e$: Study of a Bi-directional Jet Flap Device for Application to Helicopter Rotor Blades," Honeywell Document No.
12081-FR1, July 1971 18 Englar, R.J., williams, R.M., "Design of a Circulation I Jones, G.S., Banged, L.S., Garber, D.P., Huebner, Control Stern Plane for Submarine Applications," L.D., McKinley, R.E., Stutton, K., Swanson, R.C., NSRDC Technical Note AL-200, March 1971 Weinstein, L, "Research Opportunities in 17 Englar, R.J., williams, R.M., "Test Techniques for Advanced Aerospace Concepts", NASA/TM-2000- High Uft, Two-Dimensional Airfoils with Boundary 210547, December 2000 Layer and Circulation Control for Applications to Rotary 2 Metral, A.R., "On the Phenomenon of Fluid Veins Wing Aircraft", Canadian Aeronautics and Space and their Application, the Coanda Effect," AF Journal, Vo1.19, No.3, March 1973 Translation, F-TS-786-RE, 1939 18 Anderson, W. K., and Bonhaus, D. L., "An Implicit 3 Schlichting, H., "Boundary_ Layer Theory", page Upwind Algorithm for Computing Turbulent Flows on 750 - 751, 2 '< Edition, 1979 Unstructured Grids," Computers Fluids, Vol. 23, No. 1, 4 Karamcheti,K., "Princioles of Ideal-Fluid 1994, pp. 1-21.
Aerodynamics", John Wiley and Sons, Inc., page ,8 Spalart, P. R., and Allmaras, S. R., "A One-Equation 376 - 401, 1966 Turbulence Model for Aerodynamic Flows," AIAA s Wood, N., and J. Nielson, "Circulation Control Paper 92-0439, January 1992.
Airfoils Past, Present, and Future," AtAA Paper 2o Marcum, D. L., "Generation of Unstructured Grids for 850204, January, 1985.
Viscous Flow Applications," AIAA Paper 95-0212, 6 Englar, R.J., Circulation control Pneumatic January 1995.
Aerodynamics: blown force and Moment 2, Marcum, D. L, and Weatherhill, N. P., "Unstructured Augmentation and Modifications; Past, Present, & Grid Generation Using Iterative Point Insertion and Future", AIAA 2000-2541, June 2000 Local Reconnection," AIAA Journal, Vol. 33, No. 9, 7 Liu, Y., Sankar, L.N., Englar, R.J., Ahuja, K.K., September 1995.
"Numerical Simulations of the Steady and Wentz, W.H., Seetharam, H.C., "Development of a Unsteady Aerodynamic Characteristics of a Fowler Flap System for a High Performance General Circulation control Wing Airfoil," AIAA 2001-0704, Aviation Airfoil," NASA CR 2443, 1974 January 2001 z3 Braden,J.A., Whipkey, R.R., Jones, G.S., Litley, D.E., 8 Smith, A.M.O., "High Lift Aerodynamics, 37 t" "Experimental Study of the Separating Confluent Wright Brothers Lecture, AIAA 74-939, August Boundary Layer," NASA CR 3655, June 1983 24 Englar, R.J., =Low-Speed Aerodynamic 9 Anderson W.K., Bonhaus D.L., McGhee R.J., Characteristics of a Small Fixed Trailing Edge Walker B.S., "Navier-Stokes computations and Circulation Control Wing Configuration Fitted to a experimental comparisons for multi-element airfoil Supercritical Airfoil," DTNSRDC Report ASED-81/08, configurations." J Aircraft 1995;32(6):1246-53.
March 1981.
10 Englar, R.J., Circulation control - A Bibliography van Dam, C.P., =lhe Aerodynamic design of Multi- of DTNSRDC Research and selected Outside Element High-Lift Systems for Transport Airplanes," References., DTNSRDC-84/052, Sept 1984.
Progress in Aerospace Sciences 28 (2002) 101-144 11Englar, R. J., "Development of the A-6 Circulation Schaeffler, N.W., Hepner, T.E., Jones, G.S., Control Wing Flight Demonstrator Configuration," Kegerise, M.A., "Overview of Active Flow Control DTNSRDC Report ASED-79/01, January, 1979.
Actuator Development at NASA Langley Research 12Oyler, T.E., Palmer, W.E., "Exploratory Center," AIAA 2002-3159, June 2002 Investigation of Pulse Blowing for Boundary Layer Control", North American Rockwell Report NR72H- 12, January 15, 1972 ,3 Waiters, R.E., Myer, D.P., & Holt, D.J., "Circulation Control by Steady and Pulsed Blowing for a Cambered Elliptical Airfoil", West Virginia University, Aerospace Engineering TR-32, July 14Cagle, C.M., Jones, G.S. "A Wind Tunnel Model to Explore Unsteady Circulation Control for General Aviation Applications, AIAA 2002-3240, June 2002 AIAA 1 = Flow Control Conference 18 AIAA 2002-3157