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AIAA 2002-6006
Flow Control Research at NASA Langley in Support of High-Lift Augmentation
William L. Sellers III, Gregory S.Jones, and Mark D. Moore NASA Langley Research Center Hampton, VA
Biennial International Powered Lift Conference 5-7 November 2002 Williamsburg, VA
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FLOW CONTROL RESEARCH AT NASA LANGLEY IN SUPPORT OF HIGH-LIFT AUGMENTATION
William L. Sellers III*, Gregory S. Jones t and Mark D. Moore _ NASA Langley Research Center Hampton, VA 23681-219
Abstract M Mach number The paper describes the efforts at NASA ,, Mass flow (lb/sec) Langley to apply active and passive flow control U Velocity (ft/sec) techniques for improved high-lift systems, and q Dynamic Pressure (lb/ft 2) advanced vehicle concepts utilizing powered w Slot Width (inches) x Streamwise distance high-lift techniques. The development of simpli- y Vertical distance fied high-lift systems ,utilizing active flow z Spanwise distance control is shown to provide significant weight Xt_ Distance from actuator to TE and drag reduction benefits based on system X_v Distance from actuator to separation point studies. Active flow control that focuses on a Angle of attack (degrees) separation, and the development of advanced 6s Slat deflection (degrees) circulation control wings (CCW) utilizing f Frequency (Hz) unsteady excitation techniques will be dis- Rec Reynolds number based on cbord length cussed. The advanced CCW airfoils can provide U Mean streamwise velocity multifunctional controls throughout the flight envelope. Computational and experimental Abbreviations data are shown to illustrate the benefits and issues with implementation of the technology. AFC Active flow control Nomenclature LE Leading edge b Span (inches) TE Trailing edge Ct 2-D sectional lift coefficient CCW Circulation Controlled Wing
CL 3-D Wing lift coefficient Superscripts Cdpu 2-D sectional profile drag coefficient Cp Pressure Coefficient ' Fluctuating component C Airfoil chord (inches)
% Steady blowing momentum coefficient, = J/cbq Subscripts
"< C_> Oscillatory blowing momentum coefficient, = jet Conditions of jet
Cla = (0.10; 0.03)%. The LaRC data was at 0.8 C Rc=37.6x106, M=0.3, and Cg = (0.27; 0.05)%. The t
results demonstrated that when normalized by Cim_x, 0.6 *=0.0, R¢=O.0M" there was no significant Re dependence on the lift bene- fits with separation control. 0.4 p /_ _B..F.__O.0, Rc=28"2M"
0.2 tt +"F*=O.7, Rc=8.0M" 1.2 , , ' ' i , ' ' ' i .... i , , , , i , , , ' 0 _ -m -"F'=O.7, Rc=28.2M"
.0._} .... i .... i .... n .... i .... i .... i .... i .... 1 0.01 0.02 0,03 0.04 0.05 0.06 0.07 0.08 0,09 C ¢lpu
0.8 Fig. 3 Lift-form drag polars for fir = 20% M=0.2, and C7(0.0; 0.05)%
CI 0.6
1.3 ' ' ' ,,,,,i , , , ,,,,,i , , , ,,,,,i , , , ,,,,, Steady Blowing (_ 1.2 0.4 /t_ --El- F+ = 0.0 Tau -m- Periodic Blowing / ,4g .L /_ _ F+ = 2.1 LaRC 1.1 O.2 rTl F+ = 0.5 --m- F+_- 2.6 TAU 1 01 ...... C I 0 5 10 15 20 25 0.9 c_,deg 0.8 •
The benefits included an increase in c_ at Clm_x of 2 ° and 0.6 a milder stall than the baseline airfoil. In the case of the
flap, the data in figure 3 shows that F + and C_ were not 0.5 affected by Reynolds number. It also shows the signifi- 0.001 0.Ol o.1 1 10 cant increase in lift for a fixed a = 4.0 °. An F+=0.7 was c,% effective in fully reattaching the flap at low angles of attack, but the effectiveness decreased at higher angles Fig. 4 Comparison between steady and unsteady of attack. This was caused by F + being tuned to the blowing for flap separation control lower co's, and insufficient % for the higher angles. Perhaps the most important finding from this investiga- The results from this investigation were so tion was the fact that oscillatory blowing was more than promising that NASA LaRC contracted with the Boeing two orders of magnitude more effective than steady Company to put together a team to conduct a system blowing. This is illustrated in figure 4, where for a Cl _- study _3 of oscillatory blowing for separation control. 1.0 oscillatory excitation required a momentum coeffi- They examined a number of potential applications, and cient of only 0.03% compared with 2.7% for steady identified high-lift separation control as the highest blowing priority. The study team relied on unsteady separation control to only match the performance of a conven- tional high lift system. The major benefits would there- fore come from developing a simpler, lighter, and less costly system based on flow, control. Their concept of a simplified high-lift system evolved to a system with a simple hinged drooped leading edge and simple hinged detrimental effect on airfoil performance, due to the trailing edge flaps. Their estimates of potential benefits creation of a second shock wave at the blowing slot. were based on applying active flow control to both LE When excitation was applied slightly ahead of the and TE devices. Their benefit assessment demonstrated shock it proved beneficial regardless of the Maeh potential reductions including a 2.6% in part-card number. count, 3.3 % in empty weight, and 1.3% in manufac- turing cost. These benefits combined to provide a 3.3% Further investigations 8'9'_°of compressibility and reduction in drag due to a simpler and lighter high-lift sweep effects were conducted with a new "hump" system. The system study provided several recommen- model. The model derives its name from the fact that it dations as to future research based on their experience uses the coordinates of the upper surface of a Glauret in the study. A high-lift system utilizing active flow Glas lI airfoil and is mounted on the tunnel sidewall to control would require flap deflections between 50 ° and form a "hump" as shown in figure 6. The experiments 60 °. That was beyond the range of all available data were conducted in the LaRC 0.3-meter Transonic Cryo therefore, it was recommended that additional research Tunnel 14'15(TCT). be focused in that range. In addition, no data was available for active separation control with drooped 0.30 • • • i • • , i , , • i ...... LE's devices. Lastly the team recommended that a 0.25 ......
study of separation control with simultaneous LE and 0.20 ...... TE blowing was needed to determine if there were x/c = 0.59. adverse interactions. y1_o,_5...... _...... +....._.._...... At NASA LaRC there were concerns regarding compressibility and sweep effects, and excitation mode ...... i ...... ""...... i-- that needed to be addressed, since all could become 0.00 0.20 0.40 0.60 080 t .00 issues in a viable high-lift system. A limited investiga- x/c tion 7 of compressibility and dynamic effects utilized the earlier 0015 model with LE blowing to assess Fig. 6 Hump model geometry and blowing slot oscillatory excitation in the presence of shocks, and to locations examine dynamic effects in the wake. The investiga- The original location of the wing leading edge is tion showed that similar to low Re experiments with designated as x/c=0.0 and the region between -0.05 < incompressible flow, Cl.... increased by 15%, post stall x/c < 0.05 was faired smoothly into the tunnel floor to lift increased by 50%, and drag was reduced by 50% in avoid separation at the leading edge. This smoothing compressible flows. Figure 5Fig. 5 shows typical had a dual effect. First it prevented separation at the results comparing the lift from the Mach 0.28 and 0.4 wing leading edge, and second it meant that the model cases. would be fully turbulent due to the upstream turbulent wall boundary layer. The airfoil thickness was 20% of 1.6 = = i i i = the root chord and without flow control the flow "-'A'-C¢ F'=O, M=0.28 separates at x/c=0.65 both with and without sweep. 1.4 -'k--c1, F*=2, M=0.28 Oscillatory excitation slots were placed at x/c=0.55 and
l"2 "{3-c,' F'=O' M=0'4 _ x/c=0.64. 1.0 , =. , =. i0% The results of this investigation showed that C steady suction or blowing with a momentum coefficient 0.8 ranging from 2% to 4% was required to fully reattach
0.6 the flow and recover the predicted ideal pressure distribution. Oscillatory excitation was comparable to 04 steady suction in effectiveness and significantly more
02 effective than steady blowing. Although periodic excitation was capable of reattaching the flow, it is not 0.0 , , , ,, , , , capable of reproducing the same pressure jump that 2 4 6 8 10 12 14 steady suction produces across the excitation slot. (x, deg Figure 7 shows that the superposition of a slight amount Fig. 5 The effects of Mach number and excitation on of steady suction with periodic excitation enhances the the lift of the NACA 0015 airfoil, Re = 12.7x106, <%> receptivity of the shear layer and produces better = 0.03% (M=0.28) and 0.0245% (M=0.4) results. The results were sensitive to the excitation location. If excitation was well ahead of the shock location it had a -1 T T r was equipped with a 15% chord simple-hinged drooped -O-Baseline LE, which will be referred to as a leading edge slat and -o.8 =_ "-_-c =-0.036% a 25% chord simple-hinged trailing edge flap. A -o.6 -e-c = 0.002% comparison between the cruise configuration data with results taken previously at the NASA Langley Low ..0.4 Turbulence Pressure Tunnel (LTPT) was very good. c The model was constructed in a modular fashion p -o.2 that provided the insertion of flow control actuators at several locations as shown in figure 9. The LE and o trailing edge flap settings were all computer controlled. The LE slat could be deflected to -30°and the trailing 0.2 edge flap could be deflected to -60 °. A photograph of the model installed in the NASA Langley Basic Aero- 0.4 dynamics Research Tunnel (BART) with the LE and -tl.5 0 0.5 1 1.5 TE flaps deflected is shown in figure 10.
)OtC
Fig. 7 The effect of steady blowing and suction on periodic excitation, F÷=I.15, <%>=0.23%, Rc=16xl06, M=0.25
Sweep and additional compressibility data were documented J° in a follow-on investigation. The results shown in figure 8, illustrate that compressibility tends Sill to elongate the separation bubble due to reduced mixing above the separated region. Periodic excitation is capable in reducing the separation bubble at compressi- Fig. 9 Modular EET model used for experiment, ble speeds however; its effect is reduced when using the c=406.4mm frequencies and excitation levels used at incompressible speeds.
-1.5 r • . r i =l -_-Baseline, M = 0.25
/ "_ -E}-M=0.25,
Fig. 10 Simplified high-lift version of EET airfoil model installed in BART
0.5 ...... , ...... I ...... The results using LE slat flow control was 0.5 1 1.5 x/c published in reference 12. The slat actuator had two blowing slots located at x/c = 0.14 and x/c = 0.3. The Fig. 8 Comparison of control effectiveness at forward slot was hidden under the main element at the compressible speeds, M=0.65, l_=30x106, F+=0.3, stowed condition to eliminate any actuator cavity inter- and M=0.25, IL:=16xl06, F+=0.4 actions with the freestream. A well-optimized airfoil might suffer from discontinuities caused by blowing An investigation of the use of oscillatory blow- slots, and the results showed that up to C_max,there was ing based on a modern supercritical wing has begun to no measurable effect. An investigation of the baseline answer the remaining questions from the Boeing (eg. no AFC) performance of the LE slat was as systems study. The NASA Energy Efficient Transport expected. The results in figure 11 show that slat (EET) airfoil la was chosen as the baseline. The model deflection increased maximum lift by 12%. The angle atwhichstalloccursis delayed by approximately 4 degrees. The effect of AFC was to increase the lift by Additional testing is underway at NASA LaRC another 12%. documenting the effect of oscillatory blowing on TE 2.011 flap separation, and combined LE and TE blowing. -e-Baseline ,_ [ Plans are underway to construct a larger model for 1.95 -_-Controlled testing in the LTPT tunnel. This facility is designed 1.90 specifically for 2-D airfoil testing and the larger size
1.85 will remove some of the current limitations previously encountered with the actuator placement in such small Ci, r_mx1,80 elements.
1.75 Circulation Control 1.70 The aerodynamic characteristics of Circulation 1.65 Control Wings (CCW) have been experimentally and
1.60 I numerically studied. _7' _,_9.20for more than 65 years. -5 -1'0 -i'5 -2=0 -25 -30 The benefits of CCW have also been highlighted in _)s' deg flight tests that focus on high-lift. In spite of significant improvements in performance, CCW techniques have Fig. 11 The effect of flap shoulder flow control on lift not been applied to production aircraft. Many of the roadblocks to this technique include engine bleed During the test different types of excitation were requirements, and cruise performance penalties. LaRC investigated. Previously actuators were typically driven has three current research efforts that focus on devel- by a continuous sine wave at varying frequencies and oping CCW technologies, 1) pulsed pneumatic flaps for amplitudes. The actuation frequency is typically the general aviation, 2) pneumatic diffusers for distributed resonant frequency of the piezoelectric diaphragm and engines, and 3) channel wing. in this case was either 853 Hz or 1 KHz depending on Traditionally Circulation Control Wings (CCW) the type of piezo element installed. It is a balancing act are restricted to a pneumatic modification of the flow trying to match the actuator frequency with the field through a Coanda effect 2_ as shown in figure 13. optimum F_ for separation control. In this test the The wall bounded jet flows along the surface and has actuator was also driven using an amplitude modulation the nature of a boundary layer near the wall but (AM). This technique allows you to operate the piezo becomes that of a free jet at a larger distance from the actuator at its optimum frequency and tune the separa- wall. 22 The degree of jet turning can be related to the tion control to appropriate F+ using the AM slot height, surface radius, jet velocity, and the Coanda modulation. Figure 12 shows a comparison of the surface geomet_. effect of AM modulation for different values of
2.7 , , ,
2.6 []
f" 2.5 , / • ¢ I 2.4 ,'[2
2.3 _'"' Fig. 13 Example of Coanda effect
2.2 El/'_,,,_ --I_ F*=I 2. 2 Appling the Coanda effect to the trailing edge of _=12.2, F+AM----0.74 an airfoil results in streamline turning similar to that of 2.1 [] _ J 0 0.004 0.008 0.012 0.016 a traditional flap/slat high lift system. Figure 14 shows
UPPER STEADY ACTUATOR UPPER MANIFOLD DIFFUSER SLOT
LOWER STEADY t PULSED / MANIFOLD \ ACTUATOR LOWER ACTUATOR SLOT MANIFOLD
Fig. 16 Two-Dimensional 17% Supercritical General Fig. 14 CFD simulation of streamline turning Aviation Circulation Controlled Airfoil with a Both of these parameters are dependant on the jet circular trailing edge, r/c=2% velocity ratio and Coanda surface and the airfoil leading The optimization of high lift and cruise edge geometries. The resultant lift performance char- performance with one airfoil shape gives rise to the acteristics also highlight the limits of separation control pneumatic flap concept 24'='sThis concept is based on the and the effect of super-circulation as shown in figure ability to switch from a high lift configuration to a 15. cruise configuration without utilizing any mechanical systems. Having two independent blowing systems 4.0 ,., _.1_. ,..i .... 1 .... = .... i .... i .... t .... allows one to have such a multi-function system that can be used for high lift systems and flight control systems such as ailerons and air brakes. Moving from a 3.0 Sep arat°n _I f high lift to a cruise configuration is dependent on the Control -I _e_ upper and lower blowing ratios and the free stream velocity, and can be optimized for minimum cruise drag. Ci 2.0 _c c- z8 It was important to establish a baseline using steady blowing to determine the effectiveness of pulsed circulation control. The effectiveness of the pulsed pneumatics is referenced to the steady blown CC 1.o _c/_c.so I conditions at ct = 0.0 degrees. The baseline i Super-Circulation.__ performance of the GACC model is characterized by Control t the lift as shown in figure 17. The GACC experimental 0.0 .... ' .... ' .... i .... , ...... , ...... lift results evaluated at an angle of attack of zero will 0.00 0.02 0.04 0.06 0.08 provide lift augmentation, ACL/AC_,=50. This is Cmu consistent with other small trailing edge CCW airfoil experiments. The test matrix was limited to lift coeffi- Fig. 15 CCW Lift augmentation through separation cients of approximately 3. This was consistent with and super-circulation flow control for GACC airfoil general aviation requirements for an aircraft with a AFC Pneumatic Flap (GACC) wing sized for cruise conditions and still be able to The LaRC General Aviation Circulation Control meet FAR stall requirements. It also limited the wall (GACC) Wing research effort is intended to address interference of the GACC model in the BART tunnel technology issues, such as scaling, mass flow, and noise requirements. A 2-D flow physics supercriticat airfoil model :3 (figure 16) with dual slotted circulation control capability has been designed and built for low speed testing in the BART Tunnel. The would resemble the steady state blowing series. This TARGET I would assume a perfect square wave response at the jet PERFORMANCE II 0 exit. The reality of a perfect square wave diminishes 3.0 f...... _ ___ _ u__C__...... with the complexities of the actuator system. The response of the state-of-the-art high-speed actuator valves used for this study does not generate a
C, 2.0 t" J _.,_ "-E_0.04] perfect pulsed 26 waveform. Transmitting the pulse F ,_ ...w, _ -_o.ozs through the nozzle and into the nozzle exit distorts the F ,._ L,_'_-_,w_H_I -*-o.o14 waveform as shown with thin film data located at the " _ _,-_,_" ---0.007 nozzle exit. For the low frequency pulsed jet, the effect "°I JZf- ---o.ooo of duty cycle is shown in figure 19.
0 0 ...... i , , , , , .... , . • , . 1.2 " " Duty Cycle' "
" 1 0 -S 0 S I0 IS 0
Lncrel)kn$* Duty C?cle c_, deg _1.0 e?-...... 5.... Fig. 17 Lift characteristics of the 2-D GACC airfoil o >0.8 using steady blowing E
LL Dual Blowing ._ 0.6 .= To minimize the drag the GACC airfoil is equipped with upper and lower blowing on the Coanda _o.4 N surface. In principle the pneumatic flap is optimized _O.Z for drag when the upper and lower Coanda jets close O the wake with minimum momentum losses. This Z
0.0
occurs when the upper and lower jets remain tangential. 0.00 0.02 0.04 0.06 0.08 0.10 If the two jets turn around the Coanda surface and Time, sec impact upon one another the flow will rebound creating Fig. 19 Normalized thin-film time history for pulsed a larger wake and increase in drag. This occurs at jet CCW at the slot exit, h=0.020, and driver velocity ratios above 2. The trends in dual blowing frequency=35 Hz efficiency, shown in figure 18, demonstrate a potential 20% improvement in L/D at an AOA of zero degrees. The peak amplitude for the low duty cycle conditions (20% and 30%) does not reach the maximum output 40.0 performance of the actuator system. The actuator valve i U /U ) being closed before the plenum and actuator volumes JET 0 I 30.0 _0.0 ' .._ _F L_L _-i have had time to be fully pressurized causes this result. -_ o. 68 .____ Once the valve is given a close command the plenum 2o. o -_ I. I 8 /_ I____.__-_-_k'_L- remains pressurized and continue to bleed air through 39 -e z. /-/Sf_ll _v_ _-_-'_-'_L the jet exit until the plenum pressure reaches ambient L,o - 273 conditions.
I0.0 11 As the drive frequency is increased, figure 20 i shows that the rise time or valve opening distortions 0.0 increase. For the closed portion of the duty cycle, air continues to bleed from the plenum until the open I -10.0 ...... ' ...... command is given, resulting in the jet velocity not 0 -5 0 S 0 going to zero. This process limits the mass flow to the c_deg jet exit as indicated by an overall reduction in the peak Fig. 18 Dual blowing Lift to Drag ratio (h=0.01) velocity. In spite of the limitations of the actuator system, the peak velocities do approach sonic Pulsed Blowing conditions. The effectiveness of pulsed blowing on the perform- ance of the GACC airfoil is dependent on the efficiency of the actuator system. This system must include the actuator performance, diffuser performance, and the response of the internal volurne prior to the jet exit as well as the external time dependent Coanda effective- ness. Ideally the time dependent Coanda response 1.2 system concepts (RASC) and architectures to identify
¢J Increasing new mission approaches, and associated technologies .- - p ...... Duty Cycle ...... 1.0 Decreased (40% < DC < 80%) that enable these missions to be implemented. This • , )'item Elfi¢iency effort includes experts from the various field centers, EO,8 . kq academia, and industry. Although RASC conducts many studies across the NASA Enterprises, the _o.G- 'I,!i Personal Air Vehicle Exploration (PAVE) study has P- ( [I direct application to powered lift research and configu- i¸¸¸¸i,cl rations. The objective of PAVE is to explore future !i il" mobility requirements aimed toward a system solution 0.0 I !. . t_ . for door-to-door personal travel. The improvements in 0.02 0.00 0.04 0.06 0.08 0.10 lifestyle for the traveling public would be significant as Time, sec highlighted in figure 22. Small companies have tried to Fig. 20 Normalized thin-film time histories for develop personal air vehicles (PAVs) that blend the pulsed CCW at the slot exit, h=0.020, and driver functions of an automobile and an airplane for more frequency = 100 Hz many years, but typically with minimal investments in These data are consistent with conditions measured for advanced technologies to solve the inherent problems. They proposed or built various vehicle concepts that both slot configurations. For low frequency pulsed jets, would either break apart, fold, or were Vertical Takeoff the frequency characteristics of the jet can be separated and Landing (VTOL) concepts. into pulsed and turbulence regimes. The transition
from the pulsed to the turbulence regime seemed to be 400 independent on the jet velocity and occurred near 300 Hz. t...... i...... Comparing the pulsed and steady lift perform- -= 300 I- ...... i!Fast...... PAV'al _ ...... mllmm::i...... i...... ance of the GACC airfoil, figure 21, a distinct improvement in lift or mass flow required can be seen. For a given lift coefficient of 1.0, a 48% reduction in u_ ...... r-!--_- ...... lk.._ ..... 4-"_ ...... mass flow is realized for a 20% duty cycle. As the duty 200 O
cycle is increased the performance benefit decreases. 150 ...... ; ...... • ----_ ...... ; ...... Comparing the lift performance of the pulsed and steady CCW at a fixed mass flow of 25 SCFM results o Ir f ...... i in a 35% lift improvement. , • : : :
: Highways : 0 7""- ._@_.------"s, H ig hway_ ', . , 2,0 • • ' i ' ' , i • • • i • , • 1 • • j 0 500 1,0DO 1,500 2,000 2,500 o/< Great Circle Distance, nrn oc=Bo_.._ d ,.s oc=so_ __..-'_ Fig. 22 Door to door travel time by mode NASA's PAVE project is approaching this effort differently from prior efforts by studying the PAV concept through a "systems" approach with identifica- tion of the key enabling technologies and the potential I --O- STEADY benefits as investments bring about these technologies. 0.5 -E3- PULSED These new vehicle concepts fill a matrix from Conventional Takeoff and Landing (CTOL), to Short
0.0 " " ' ' ' ' ' ' ...... ' " ' " Takeoff and Landing (STOL), to Extreme STOL, 0 20 40 60 80 O0 VTOL, and from single aerial use only concepts to mdot, SCFM dual-use (roadable) concepts. This matrix trades off Fig. 21 Comparison of pulsed and steady circulation complexity and cost with the ability to achieve a nearer control, frequency = 35 Hz, and varying duD' cycle door to door transportation system. As mentioned in the NASA Blueprint, these new PAV concepts will F.uture Applications have to be extremely safe and simple for use by non- The Langley research Center (LaRC) was professional pilots. They will require very low takeoff chartered by the NASA Administration to be the lead and landing speeds. Because of their projected wide NASA Center for evaluating revolutionary aerospace use, they must be environmentally friendly with low
10 noise and automotive equivalent emissions. Cost will out a waste gate at this condition. Thus a no cost air be a big factor in the development of PAVs as the supply is present, with air mass flow on the same order vehicles must provide a positive return on investment as are required for a moderate performing CC system. compared to an individual's value of time. Lastly the Another key concern becomes the ability to achieve vehicles must be sized to fit into the limits imposed by high lift without engine power, and thus without the existing infrastructure. Not to minimize the importance turbocharger air supply. Reservoir air plenums would or difficulties in the revolutionary changes required in require a significant weight for even the minimum the airspace system itself, the rest of this discussion will several minute capacity. However, the use of slow focus on vehicle technologies required for PAVs. burn, solid rocket gas generator offers a much lighter Part of the PAVE study objective was to explore emergency-only solution. The resulting wing weight the PAV design space, which included defining, estab- reduction more than compensates for the valves, lishing, and integrating synergistic technologies based emergency system, and ducting required for the CC on a 2015 TRL. To accomplish this task it is necessary system, providing a lighter, smaller, and more compact to develop advanced concepts utilizing physics based wing. Circulation control could potentially be highly methods, and comparing concepts to l), reference synergistic in several more exotic application areas, baselines, 2) each other, and 3) alternate travel modes. including the use of distributed propulsion systems. An important output from this approach is an identifi- Two examples of NASA PAVE concepts that utilize cation of the sensitivities, and the gaps in various advanced pneumatics in a highly synergistic fashion technologies. will be described below. At this stage in the project, the study has broken down the vehicle technologies into subsystems that Channel Wing include propulsion, aerodynamics, aero-propulsive NASA in collaboration with Georgia Tech systems and controls, and structures categories. One Research Institute is revisiting the channel wing must keep in mind that many of these concepts are concept 27"28,to develop very high lift for Extreme highly integrated and all four of these areas must be STOL (ESTOL) applications, but with a very simple worked simultaneously. Propulsion technologies system, that has no externally moving parts. The include reciprocating engines with increased specific powered-lift Channel Wing concept combines Circula- output, lightweight and low-noise turbofan and turbo- tion Control aerodynamic and propulsive technologies shaft engines, distributed propulsion mini-engine to provide a Pneumatic Channel Wing configuration systems, and electric propulsion systems. intended to have ESTOL or perhaps even near VSTOL PAV concepts vary from the mild tothe wild, capability, but without the ability to hover. The and Circulation Control (CC) has been applied across application of CC to the channel wing solves an this range because of the inherent low takeoff and important problem of the original channel wing landing speeds involved in these missions. Circulation concept. While the channel section was essentially Control is most effective at lower velocities since the stall-proof, generating increasing lift up to 45 degrees key driving parameter is the jet velocity compared to angle of attack, this lift was unusable in takeoff or freestream velocity ratio. Thus, lower takeoff and approach because of limited tail scrape angles of the landing velocities can utilize lower jet velocities or fuselage (typically about 12 degrees). The use of CC lower mass flows while achieving the same effective provides increased channel circulation lift in a large lever-arm on the maximum lift. Examination of current streamtube, at usable angles of attack. A preliminary General Aviation (GA) baseline vehicles show that the design study of this pneumatic vehicle is based on 61 knot stall rule has locked these aircraft into low wing previous wind tunnel and flight-test data. Advanced loading solutions, which are highly susceptible to gust flow control technologies are integrated into a simple concerns, and result in wings that are considerably Pneumatic Channel Wing (PCW) configuration, shown larger that that desired in cruise. As an example, in figure 23. Preliminary wind-tunnel development and current GA aircraft achieve an L/D of approximately 10 evaluations of a PCW powered model have shown to 12 at cruise, even though their L/Dm_x is about 17. substantial lift capabilities for the CCW blown channel Applying complex and expensive high lift systems as in wing configuration a_, CCMAXapproaches 8.5 to 9.0 transport aircraft is not a solution that will carry well to using just the channel portion of the wing, i.e. no out- this marketplace, or to the experience level of GA board blowing. The blown model also showed the operators. One of the simplest uses of Circulation ability to interchange thrust and drag by varying Control applies the GACC airfoil with the utilization of blowing to provide greater flexibility in Super STOL a turbocharger to power the pressurized blown plenum. takeoffs and landings. Adding the outboard CCW w/ Since turbocharging is merely used for altitude com- pulsed trailing edge pneumatics is expected to increase pensation in GA aircraft, and not increased power at the high lift performance and provide improved flight takeoff, all the turbocharger compressed air is dumped control capability.
11 ducting the flow onto a Coanda surface for pneumatic control of the diffuser. Coupling the pneumatic nacelle with the MGGF concept provides the smallest propul- sion system power requirement possible for V/STOL. The MGGF provides a relaxing of the engine-out in hover approach sizing constraint which determines the engine size, dramatically reducing the required thrust to weight of the vehicle from approximately 2.4 for a twin engine concept, to about 1.4 for MGGF concept while also eliminating the need for an engine cross-shafting system. The pneumatic nacelle provides maximum hover nacelle entrainment and lip suction for a cruise shaped nacelle in the hover sizing condition. It also provides nacelle separation control in transition and crosswind conditions. There are penalties involving a more complicated fuel routing system and fan redundancy. A recent study showed that only 4% of all propulsion system problems were related to the fan, while 71% were due to gas generators making the Fig. 23 Artist drawing of a powered lift channel wing with a CCW aircraft concept redundancy of the MGGF a potentially safer system. The combination of these two technologies makes for a Distributed Propulsion more robust and safe aircraft system. Pneumatic control and distributed engine technologies are being applied to a tilt nacelle concept based on the Grumman 698 design. The concept shown in figure 24, utilizes thrust vanes to generate all hover control authority. In addition, the concept uses pneu- matic nacelles that provide pneumatic morphing of the nacelle allowing a virtual tailoring of the inlet and exhaust flow pattern. This will enable the designer to maximize the propulsion system performance through- out the flight envelope. Additional benefits include a reduction in the effective hover disc loading at the ground plane and a reduction in the ground erosion and Fig. 25 Multi-Gas Generator Fan with Circulation foreign object damage (FOD) constraints. Control Nacelle
ConcludinR Remarks
NASA Langley has a long-standing research effort in separation control for high-lift systems that includes both passive and active approaches. The current emphasis has shifted to unsteady methods for active flow control, and the two approaches currently under development include oscillatory blowing for separation control and pulsed circulation control technology. The active separation control research has taken the technology from low Reynolds number laboratories to Reynolds numbers representative of flight conditions. The results to date have examined both leading edge and flap separation control, Mach effects Fig. 24 Personal Air Vehicle Concept utilizing and sweep effects (e.g. infinite sweep). The lift benefits circulation control nacelles of active separation control had no significant Reynolds The Multi-Gas Generator Fan (MGGF) shown in number effects. Oscillatory excitation was shown to be figure 25, is based on a concept that uses the exhaust of two orders of magnitude more effective than steady excitation. The results were shown to be sensitive to several small engines to drive a tip driven fan, then
12 Mach effects, and understanding the flow physics was Acknowledgements essential to placing the blowing slots at the optimum location. Langley is presently attempting to extend the The authors would like to gratefully acknowl- results from the simple airfoils previously tested to a edge the outstanding collaboration between Avi Seifert modem supercritical airfoil with drooped leading edges of Tel Aviv University and LaTunia Pack of NASA and a simple trailing edge flap. The results for the Langley that provided the contributions and material for drooped leading edge look very promising but more the active separation control research for high lift work is required to address flap separation control at systems. The authors would also like to acknowledge high flap deflection angles. the program managers who support the research efforts. The performance of the pulsed Circulation The ASCOT Project is led by Long Yip, the Morphing Control blowing system realized a 50% reduction in Project is led by Anna-Maria McGowan, and the TCAT required mass flow for a given lift coefficient. This Program is led by James Pittman. result alone opens a window of opportunity for References potential advanced pneumatic systems that have been shown to exceed the performance of conventional high lift systems. Variations in the duty cycle at a given 1 "The NASA Aeronautics Blueprint - Toward frequency highlighted the controllability of the a Bold New Era of Aviation'; 3 Jun. 2002, performance with small bursts of high-speed air. http://www.aero- Continued research is necessary to quantify the overall space.nasa.gov/aero blueprint/index.html system time dependent response of the airfoil including 2 Washburn, A. E., Gorton, S. Althoff, and the leading edge and internal plenums. The GACC Anders, S. G.: "Snapshot of Active Flow airfoil has been proven to be an excellent test bed for Control Research at NASA Langley", AIAA the multi-functional circulation control study that can Paper 2002-3155, AIAA i st Flow Control operate as a high lift system, a pneumatic aileron, and a Conference, St. Louis, MO, 24-27 June 2002 high-speed air brake. Follow-on testing is expected to 3 Gad E1 Hak: "Flow Control: Passive, Active, improve the database for CFD validation and the under- and Reactive Flow Management", Cambridge standing of the flow physics related to circulation University Press, 2000 control concepts. 4 Lin, John C.: "Control of Turbulent Boundary The tight coupling of advanced flow control Layer Separation using Micro- Vortex Genera- and propulsion systems is an inevitable solution for tors", AIAA Paper 99-3404, 30th AIAA Fluid vehicles with optimum "system" performance. The Dynamics Conference, Norfolk, VA, 28 June- ability to impact global objectives, such as enhanced mobility that permits travel further, faster, anywhere, 1 July, 1999. anytime, can not be understood in terms of advanced Greenblatt, David, and Wygnanski, Israel, J.: technologies alone. For NASA, the challenge is to "The Control of Flow Separation by Periodic develop technologies that spur significant improve- Excitation", Progress in Aerospace Sciences ments. Clearly, from statistical data of travel trends and 36 (2000) 487-545 average speeds, the hub and spoke, and highway 6 Seifert, A., and Pack, L. G.: "Oscillatory systems are approaching their limits in many sections of Control of Separation at High Reynolds the market over the next 20 years. At these transition Number", AIAA Paper 98-0214, AIAA Aero- times, there is an especially great need to understand space Sciences Meeting & Exhibit, 36th, how advanced technologies can create a new solution Reno, NV, Jan. 12-15, 1998 space that opens up revolutionary, new capabilities 7 Seifert, Avi, and Pack, LaTunia G.: "Oscilla- instead of mearly trying to maintain what is daily being tory Excitation of Unsteady Compressible lost. One key parameter that helps define these new Flows over Airfoils at Flight Reynolds solution spaces in aeronautics is the speed range of Numbers", AIAA Paper 99-0925, Aerospace vehicles, that is, the ratio of cruise speed to takeoff Sciences Meeting and Exhibit, 37th, Reno, speed. Reaching new levels of this ratio will permit NV, Jan. 11-14, 1999 tremendous new mission capabilities for vehicles that 8 Seifert, Avi, and Pack, LaTunia: "Active takeoff and land in very limited spaces, and at the same Control of Separated Flows on Generic time have smaller wing structures more nearly optimum Configurations at High Reynolds Numbers", for cruise performance. Flow control research at AIAA Paper 99-3403, AIAA Fluid Dynamics NASA is the avenue for achieving radical improve- Conference, 30th, Norfolk, VA, June 28-July ments to this parameter, as well as other key metrics in efficiency, robustness, and cost. 1, 1999
13 9 Pack,LaTunia,andSeifert,Avi: "Dynamics 19 Englar, R.J., Circulation control Pneumatic of Active Separation Control at High Aerodynamics: blown force and Moment Reynolds Numbers", AIAA Paper 2000-0409, Augmentation and Modifications; Past, AIAA Aerospace Sciences Meeting and Present, & Future", AIAA 2000-2541, June Exhibit, 38th, Reno, NV, Jan. 10-13, 2000 2000. 10 Seifert, Avi, and Pack, LaTunia G.: "Sweep _o Liu, Y., Sankar, L.N., Englar, R.J., Ahuja, and Compressibility Effects on Active Separa- K.K., "Numerical Simulations of the Steady tion Control at High Reynolds Numbers", and Unsteady Aerodynamic Characteristics AIAA Paper 2000-0410, AIAA Aerospace of a Circulation control Wing Airfoil," AIAA Sciences Meeting and Exhibit, 38th, Reno, 2001-0704, January 2001 NV, Jan. 10-13, 2000 21 Metral, A.R., "On the Phenomenon of Fluid 11 Seifert, Avi, and Pack, LaTunia G.: "Separa- Veins and their Application, the Coanda tion Control at Flight Reynolds Numbers: Effect," AF Translation, F-TS-786-RE, 1939 Lessons Learned and Future Directions", 22 Schlichting, H., "Boundary Layer Theory", AIAA Paper 2000-2542, Fluids 2000 page 750 - 751, 2nd Edition, 1979 Conference and Exhibit, Denver, CO, June 23 Cagle, C.M., Jones, G.S. "A Wind Tunnel 19-22, 2000 Model to Explore Unsteady Circulation 12 Pack, LaTunia, G., Sehaeffler, Norman W., Control for General Aviation Applications, Yao, Chung-Sheng, and Seifert, Avi: "Active AIAA 2002-3240, i st AIAA Flow Control Control of Flow Separation from the Slat Conference, St. Louis, MO, 24-27 June, 2002 Shoulder of a Supercritical Airfoil", AIAA 24 Rose, R.E., Hammer, J.M., & Kizilos, A.P., Paper 2002-3156, AIAA 1st Flow Control "Feasibility Study of a Bi-directional Jet Flap Conference, St. Louis, MO, 24-27 June 2002. Device for Application to Helicopter Rotor 13 McLean, J.D., Crouch, J. D., Stoner, R. C, Blades," Honeywell Document No. 12081- Sakurai, G. E., Seidel, G. E., Feifel, W. M., FR1, July 1971 and Rush, H. M.: "Study of the Application of 25 Englar, R.J., Williams, R.M., "Design of a Separation Control by Unsteady Excitation to Circulation Control Stern Plane for Subma- Civil Transport Aircraft" NASA CR-1999- rine Applications," NSRDC Technical Note 209338, June 1999 AL-200, March 1971 14 Ladson, C. A. and Ray, E. J.: "Evolution, Schaeffler, N.W., Hepner, T.E., Jones, G.S., Calibration, and Operational Characteristics Kegerise, M.A., "Overview of Active Flow of the Two-Dimensional Test Section of the Control Actuator Development at NASA Langley 0.3-meter Transonic Cryo Tunnel", Langley Research Center," AIAA 2002-3159, NASA TP-2749, 1987. 1st AIAA Flow Control Conference, St. Louis, is Rallo, R. A., Dress, D. A., and Siegle, H. J. A.: MO, 24-27 June 2002 "Operating Envelope Charts for the Langley 2_ Pasamanick, Jerome, "Langley Full-Scale- 0.3-meter Transonic Cryogenic Wind Tunnel", Tunnel Tests of the Custer Channel Wing NASA TM-89008, 1986 Airplane," NACA RM L53A09, National t6 Lin, J. C. and Dominik, C. J.: "Parametric Advisory Committee for Aeronautics, April, Investigation of a High-Lift Airfoil at High 1953 Reynolds Numbers", Journal of Aircraft, Vol. 28 Englar, Robert J., and Campbell, Bryan A.: 33, No. 4, 1997, pp. 485-491. "Pneumatic Channel Wing Power-Lift 17 Jones, G. S., Viken, S. A. Washburn, A.E. Advanced Super-STOL Aircraft", AIAA Paper Jenkins, L. N. Cagle, C.M., "An Active Flow 2002-3275, 1st AIAA Flow Control Circulation Controlled Flap Concept for Conference, St. Louis, MO, 24-27 June, 2002 General Aviation Aircraft Applications," Englar, Robert J., and Campbell, Bryan A.: AIAA Paper 2002-3157, 1st AIAA Flow "Pneumatic Channel Wing Power-Lift Control Conference, St. Louis, MO, 24-27 Advanced Super-STOL Aircraft", AIAA Paper June, 2002 2002-3275, 1st AIAA Flow Control is Wood, N., and J. Nielson, "Circulation Conference, St. Louis, MO, 24-27 June, 2002 Control Airfoils Past, Present, and Future," AI.%& Paper 850204, January, 1985.
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