Document
High Lift C ommon R esearch Model for Wind Tunnel
Testing : A n Active Flow Control Per sp e c tive
* † ‡ § John C. Lin , Latunia P. Melton , Sally A. Viken , Ma rlyn Y. Andino , ** †† ‡‡ Mehti Koklu , Judith A. Hannon , and Veer N. Vatsa NASA Langley Res earch Center, Hampton , VA, 23681 , USA Abstract This p aper provides an overview of a r esearch and d evelopment effort sponsored by the NASA Advanced Air Transport Technology Project to achieve the required high - lift performance using active flow control (AFC) on simpl e hinged f lap s while reducing the cruise drag associated with the external mechanism s on slotted flap s of a generic modern transport aircraft. The removal of the external fairings for the Fowler flap mechanism could help to reduce drag by 3.3 count s . T he main chal lenge is to develop an AFC system that can provide the neces sary lift recovery on a simpl e hinged flap high - lift system while using the limited pneumatic power available on the aircraft. Innovat ive low - power AFC concepts will be investigated in the flap s houlder region . The A F C concepts being explored include steady blowing and unsteady blowing operat ing in the spatial and /or temporal domain. Both conventional and AFC - enable d high - lift configurations were designed for the current effort. The high - lift c onfigurations share the cruise geometry that is based on the N ASA Common Research Model, and therefore , are also open geometries . A 10% - scale H igh L ift Common Research Model (HL - CRM) is being design ed for testing at the NASA Langley Research Center 14 - by 22 - Foot Subsonic Tunnel during f iscal y ear 2018. T he overall project plan, status , HL - CRM configurations, and AFC objectives for the wind tunnel test are described .
Nomenclature C = lift coefficient L C = wing reference chord; C = local wing chord for the slat and outboard flap, and ref ref C = wing chord at the yehudi break for the inboard flap ref C = momentum coefficient μ L/D = lift to drag ratio M = freestream Mach number ∞ Re, RN = Reynolds number x, y, z = coordinates along the longitud in al axis, lateral axis, and normal axis, respectively, of the HL - CRM Δ C = lift coefficient increment L 14x22 = NASA Langley Research Center 14 - by 22 - Foot Subsonic Tunnel AATT = Advanced Air Transport Technology AFC = active flow control * Senior Research Scientist, Flow Physics and Control Branch, MS 170, AIAA Associate Fellow † Senior Research Scientist, Flow Physics and Control Branch, MS 170, AIAA Associate Fellow ‡ S enior Research Scientist, Configuration Aerodynamics Branch, MS 499, AIAA Senior Member § Research Scientist, Flow Physics and Control Branch, MS 170, AIAA Senior Member ** Research Scientist, Flow Physics and Control Branch, MS 170 †† Research Scientist, Flow Physics and Control Branch, MS 170 ‡‡ Senior Research Scientist, Computation AeroSciences Branch, MS 128, AIAA Associate Fellow APU = auxiliary power unit CAD = computer - aided design CFD = computational fluid dynamics CRM = Commo n Research Model ESP = electronically scanned pressure FY = fiscal year HL - CRM = High Lift Common Research Model IR = infrared LaRC = Langley Research Center MAC = mean aerodynamic chord PIV = Particle Image Velocimetry SCF = slat - cove filler S GF = slat - gap filler STAR = Subsonic Transport Aeroacoustic Research STEP = S panwise T raversing E lectro - P neumatic Trap Wing = Trapezoidal Wing WUSS = wing under slat surface Introduction HE sizing, economics, and safety of modern transport aircraft are strongly influence d by their high - lift
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systems. The NASA Advanced Air Transport Technology (AATT) Project is seeking to demonstrate the potential benefits of reducing the cruise drag associated with modern high - lift system s without sacrificing aerod ynamic and acoustic performance during takeoff and landing operations . One possible approach is to use 2 ,3 active flow control (AFC) to provide the required lift performance while reducing the cruise drag associated with 4 , 5 the external mechanism s used to dep loy a slotted flap during high - lift operations. NASA is seeking to perform wind tunnel testing of AFC - enabled high - lift systems within the framework of the AATT Project goals to reduce fuel burn and noise of modern civil transport aircraft.
A recent system integration study indicated that up to a 2.25% fuel burn reduction is possible if an AFC - enabled simplified high lift system (i.e., simpl e hinged flaps inboard and outboard) could provide the necessary lift recovery at the approach angle of attack. The AFC - related performance gains are primarily due to the 3.3 - count excrescence drag reduction from the removal of the external fairings for the Fowler flap mechanism (see Fig. 1 (a) for a typical example ) . However, the main challenge here is to devel op an AFC system that can provide the neces sary lift recovery for a simpl e hinged flap high - lift system (Fig. 1(b)) while using the limited pneumatic power available on 7 - 9 10 , 1 1 the aircraft. Innovative low - power AFC concepts , such as the traverse actuat or and fluidic oscillators , will be investigated around the flap shoulder region . The AFC concepts include steady and unsteady blowing that operate in the spatial and/or temporal domain s . The AFC concepts for drag reduction will also leverage the knowledge 1 2 , 1 3 1 4 , 1 5 gained from the existing AFC - enhanced vertical ta il test data set , comprised of sub scale and full - scale wind 1 6 tunnel test data, and flight demonstration data , which were the culmination of several years of research and 1 7 development.
(a) A n example of e xternal fairings for Fowler (b) A simpl e hinged flap high - lift wing with AFC (no flap mechanism. external fairings) .
Figure 1. Concept of AFC - enabled high - lift system for drag reduction .
Although the initial motivation to develop an “open” high - lift geometry was for AFC research, there are also strong desires for such a geometry from government, industry, and academia for R&D efforts related to noise reduction, high - lift aerodynamics/flow physics, and CFD development/validation. For exampl e, the Trapezoidal 1 8 , 1 9 20 , 2 1 Wing (Trap Wing) model and the Subsonic Transport Aeroacoustic Research ( STAR ) model were developed to provide semispan hi gh - lift testing data for aerodynamic and aeroacoustic research. However , the Trap Wing is a generic high - lift geometry that lacks the high - fidelity details typical of a transpor t aircraft, and the aerodynamic data 2 1 of the STAR model are considered proprietary. Consequently, a new open high - lift geometry representative of a modern transport aircraft (i.e., containing relevant flow physics, aerodynamic, and aeroacoustic features) is highly desirable.
The current high - lift research effort involves both conventional (baseline) and AFC - enabled high - lift 22 , 2 3 configurations that are based on the NASA Common Research M odel (CRM). A 10% - scale high - lift model is being designed for wind tunnel testing during f iscal y ear (FY) 2018. This paper will describe the high - lift configurations and the AFC concepts currently under development for the wind tunnel test, as well as the project plan, schedule, and model design features and status.
HL - CRM Geometry Conventional H igh - L ift C onfiguration ( B aseline) 2 4 The high - lift configuration, as reported by Lacy and Sclafani, shares the CRM cruise geometry . L ikewise , the current hi gh - lift geometry is intended to be open as well. Because the original CRM cruise geometry was design ed for transonic speeds, a minor modification on the leading edge was needed to resolve high - lift aerodynamic issues at low speeds . Figure 2 illustrates t he modification for the HL - CRM configuration, as the e ffective leading - edge radius was increased to resolve a stall issue. In addition, a o ne - piece wing loft was created for easier CFD implementation.
2 4 Figure 2 . Leading edge modification of HL - CRM con figuration ( Lacy & Sclafani ) .
The conventional high - lift configuration, as predicted by OVERFLOW and CFD++, is expected to have a 2 4 maximum lift coefficient of around 2.3 to 2.4 for landing and 2.0 to 2.1 for takeoff at Re = 24.6 million and M = ∞ 0.2 (s ee Fig. 3). From the AFC per spec tive, the conventional high - lift configuration serves as the baseline that the AFC - enabled high lift configuration (see next Section) seeks to match . For example, at the landing approach angle of attack (AOA) of 8°, the li ft coefficient of the former is approximately 1.8, which is the lift value that the latter needs to recover.
Currently, CAD files of a simpler version of the CRM high - lift geometry (without the engine nacelle, landing 2 5 gear, slat and flap supporting bracket s, and horizontal tail , etc.) are on the NASA website. Th is geometry will also be used for the Geometry and Mesh Generation Workshop - 1 (GMGW - 1) at the AIAA AVIATION 2017 Forum .
2 4 Figure 3 . Lift performance of conventional HL - CRM configuration ( Lacy & Sclafani ) .
AFC - E nabled H igh - L ift C onfiguration A simpl e hinged flap high - lift configuration will be used for AFC implementation. The slat and forward portion of the main wing (i.e., less than approximately 70% of cruise chord) of the AFC - enable d high - lift configuration is the same as that of the conventional high - lift geometry . The initial AFC focus is to achieve the necessary lift recovery at an approach angle of attack (i.e., ~8°) for a simpl e hinged flap high - lift system while using the avai lable pneumatic power on the aircraft during landing when the engine is at idle power . By keeping the conventional slat, the expectation is that the maximum lift should not change significantly.
AFC Challenges The current effort on the AFC - enabled high - l ift wing leverages the knowledge gained from the recent 1 2 - 1 7 successful demonstration of the AFC - enhanced vertical tail technology . The AFC - enabled high - lift wing in the landing configuration encounters a more difficult lift recovery challenge than the AFC - enhanced vertical tail application , as summarized in Table 1. Because the lifting surface is smaller due to the elimination of the Fowler - flap system, the lift coefficient increase ( Δ C ) for the AFC - enabled high - lift system is about twice that required for L the AFC - enabled vertical tail ¾ i.e., Δ C = 0.4 4 lift increase versus equivalent Δ C ~ 0.2 ( side force enhancement L L for the latter ).
High - lift wings also generate h igher adverse p ressure gradient s compared to the vertical tail because of higher m aximum flap deflection s (i.e., ≥ 50° instead of 30° ) and h igher AOA (i.e., 8° and 16° instead of 0° and 7.5°) for aerodynamic optimization . The higher freestream speed required for the high - lift application –– M = 0.2 ( ~ 13 0 ∞ knots) versus of M = 0.15 ( ~ 100 knots) –– also reduce s the effective momentum coefficient ( C ) at the same ∞ μ pneumatic power setting. In addition, t he e ngine is at idl ing power during landing , therefore , the bleed air fo r high - lift wings is limited. T he available mass flow and pressure are moderately less than what the auxiliary power unit ( APU ) was able to provide for the AFC - enhanced vertical tail implementation . In summary, for the landing configuration , the AFC - enab led high - lift wing is required to achieve more lift with less pneumatic power when compar ed to the AFC - enhanced vertical tail.
Table 1. Challenges of an AFC - enabled high - lift system as compared to an AFC - enhanced vertical tail.
AFC - Enabled Increased Challenges for AFC - Enhanced Simpl e Hinged Flap Simpl e Hinged Flap High Lift 1 2 - 1 7 Vertical Tail High Lift Wing Wing Smaller lifting surface due to Lift Increments ~0.2 0.44 elimination of Fowler flap ( Δ C ) L system Maximum Flap Increa sed adverse pressure 30° ≥ 50° Deflection gradient AOA for 8° (approach) and 16° Increased adverse pressure Aerodynamic 0° and 7.5° (maximum lift) gradient Optimization Freestream Speed M = 0.15 ( ~ 100 knots) M = 0.2 ( ~ 13 0 knots) Effective C is reduced μ for AFC Design Pneum atic power Engine air bleed at idle Available mass flow and APU on during takeoff available power during landing pressure are reduced A recent CFD study has shown that traverse actuation can achieve ~0.2 Δ C on an AFC - enabled high - lift L system with the av ailable pneumatic power, but an additional ~0.2 4 Δ C lift enhancement is still required. The L traverse actuation system employs a small and fast moving jet packet that t ravels periodically in the span wise direction toward the fuselage. T raverse actuation at 10 Hz applied only on the inboard flap could reduce the mass flow rate requirements by about an order of magnitude compare d to nontravers e (steady) blowing, while still being able to reduce separated flow over both the inboard and outboard trailing - edg e flaps .
The current AFC - enabled high lift configuration will have a simpl e hinged flap with AFC on the flap shoulder for lift recovery on the landing configuration as described above . The model design will also have provisions for some localized AFC con cepts aimed at increas ing the lift to drag ratio ( L/D ) for the takeoff configuration . G ar ner et 2 6 al. indicated that a 1% increase in L/D for takeoff is equivalent to 2800 pound s increase in payload or 150 n autical m iles increase in range. From the fuel burn saving s perspective, the L/D improvement for takeoff could lead to a transport aircraft with smaller engines that use less fuel.
AFC S trategies The strategies for AFC - enabled lift recovery are to perform parallel investigations both experimentally a nd computationally for risk reduction purpose s . Experimentally, there are ongoing efforts to examine multiple rows of AFC actuation on the flap shoulder region o f a geometry derived from the simpl e hinged flap HL - CRM design.
2 7 , 2 8 There is some experimental evi dence from the work of DeSalvo et al. suggest ing that multiple rows of AFC actuation on a simpl e hinged flap high - lift airfoil could be effective at lift enhancement while keeping the momentum coefficient at relatively low levels.
The sweeping jet a ctuator design to be used for the current effort is similar to the Mod 2 geometry as reported 2 9 by Melton et al. U sing an array of high - speed valves, a functional traverse actuation method referred to as “ S panwise T raversing E lectr o - P neumatic (STEP) actua tors” is also being developed at the NASA Langley Research Center (LaRC). AFC approaches using multiple rows of actuators and STEP actuators will be tested in low - speed wind tunnels at NASA La RC prior to the NASA LaRC 14 - by 22 - Foot Subsonic Tunnel ( 14x22 ) test entry . Table 2 summarizes the planned AFC strategies for the HL - CRM testing at the 14x22.
Table 2. Summary of AFC test plan for HL - CRM landing configuration at the 14x22 .
Mass flow range Pressure ratio range AFC parameters Simpl e hinged flap N/A N/A N/A baseline (AFC off) Discrete nozzle Multiple row actuation and Up to 1 lb/s 1 to 3 actuators actuator spacing Multiple row actuation and Sweeping jet actuators Up to 1 lb/s 1 to 3 actuator spacing STEP a ctuators Traversing actuation Up to 1 lb/s 1 to 3 (traverse actuation) frequency and coverage Steady s lot blowing (optional — de pending Up to 1 lb/s 1 to 3 N/A on manufacturability ) N oise M easurement and R eduction H igh - lift components cause un steady flow and acoustic noise due to high - speed flow through gaps and vortical flow from component edges . From the AFC perceptive , the simpl e hinged flap should reduce the noise associated with the flow through the flap gap; however, exiting jets from AFC actuators also introduce new source s of noise.
Consequently, it is important to acquire acoustic data on an AFC - enabled high - lift system to determine whether there is a net increase or decrease in the noise levels. In addition, the acoustic portion of the HL - CRM test will 30 3 1 demonstrate two new n oise r eduction concepts: (1) flexible slat - cove filler (SCF) and (2) slat - gap filler (SGF) for airframe noise treatm ents. The goal of these concepts is to reduce the slat noise without any penalty to aerodynamic performance . The primary objective of the aeroacoustic investigation i s to t est and compare the aerodynamic performance and noise characteristics of the baseline model and the models with rigid and flexible SCF and SGF treatments. The secondary objective is to test a selected group of f lap - ed ge noise devices as reported by Khorrami et 3 2 al .
CFD Synergy There are parallel CFD efforts to explore the AFC design space at NASA LaRC (using the PowerFLOW 3 3 , 3 4 3 5 3 6 - 3 8 code for AFC and FUN3D for the baseline cases) and Boeing (using OVERFLOW ). Paramet ers to be examined may include (but are not limited to) higher flap deflection ( ≥ 50°), flaps with l onger chord s , and various aforementioned low pneumatic power AFC concepts .
Once all the model detail s are finalized after the Critical Design Review at the end of March 2017, t he final CAD files of the HL - CRM geometry for both the conventional and the AFC - enabled simpl e hinged flap cases (with 2 5 the engine nacelle, landing gear, slat and flap supporting brackets, etc.) will be uploaded on to the NASA website .
The geometry is to be used for the AIAA High - Lift Prediction Workshop s as part of its long term goals for CFD 3 9 , 40 development .
Wind Tunnel Test 4 1 The wind tunnel test will be performed at the NASA LaRC 14 - by 22 - Foot Subsonic Tunnel (14x22) , as shown in Fig. 4. The 14x 22 is an atmospheric, closed return wind tunnel with a test section 14.5 - ft high, 21.75 - ft wide, and 50 - ft long, a maximum freestream velocity of 338 ft/s, and a dynamic pressure (q) of 144 psf. The unit Reynolds number per foot ranges fr om 0 to 2.2 x10 . Test section airflow is driven by a 40 - ft diameter, 9 - bladed fan powered by a 12,000 - hp solid - state converter with synchronous motor. The tunnel has a set of flow control vanes to maintain control of the speed for low - speed testing.
A 10% - scale semispan (right wing) HL - CRM will be designed and tested in the 14x22. The aerodynamic portion of the test will be performed with closed sidewalls ( sidewall s down ) , while the aeroacoustic portion will be 3 2 performed with opened sidewalls ( sidewal l s up ) . The top insert in Fig. 1 shows an example of the latter, and the 4 2 bottom insert shows an example of the former. The landing configurations will be tested at M = 0.2, while ∞ takeoff configurations will be tested at M = 0.26.
∞ Figure 4 . NASA Langley 14 - by 22 - Foot Subsonic Tunnel . Top left insert is an example of open sidewalls for 3 2 4 2 an aeroacoustic test . B ottom left insert is an example of closed side wal ls for an aerodynamic test.
The planned experimental measurements include force s and moments using a balance (NASA MC - 110), surface pressures using pressure taps and ESP modules (for steady) and Kulite â sensors (for unsteady), model deflection using vi deogrammetry, struc ture vibration using a laser vibrometer , airframe noise using acoustic arrays (in the acoustic portion of the test), and flow visualization using tufts and infrared (IR) cameras (for transition detection). If the schedule and resource s allow, Particle Image Velocimetry (PIV) may be used to measure the off - body flow field .
Model Design The main components of the half - body model are a semispan wing, a nacelle / pylon , and a semispan fuselage with a horizontal tail. The model will primarily be used to parametrically explore the effects of different AFC systems on a simpl e hinged flap high - lift configuration with a conventional slat. The model will also be used for CFD validation and aeroacoustic measurements. Depending on the achievable mod ularity of the HL - CRM design for the 14x22, it is conceivable that this model could be used for future generic testing of advanced/ alternate AFC applications , flap/slat layouts, and aeroacoustic treatments.
Key model components such as slat, wing under s lat surface (WUSS), spoiler, and flap are all modular and replaceable. The modular approach provides flexibility and enables the model to be switch ed between conventional and AFC - enabled simpl e hinged flap high - lift configurations. The center spar remain s the same for both high - lift configurations. Provisions are also made for interchangeable model pieces for r egions at the flap and slat si de edges, wingtip, and aileron. The engine nacelle / pylon , and horizontal tail are removable on an as - needed bas i s .
A sketch of the HL - CRM in the 14x22 is shown in Fig. 5. The 10% - scale semispan model is to be installed on top of a 0.29 f oo t (~3.5 inches ) standoff (or p eniche) , and as a result , the model and its standoff will cover 68% of the tunnel span in the vertica l direction. The m odel center will be located 0.54 feet (6.5 inches ) downstream from the turntable center. The model fuselage is 20.59 feet in length; therefore, it extends past the turntable and past the end of the cart.
Figure 5 . Semi s pan HL - CRM t est in the 14x22.
Key model geometric reference parameters used for computing force and moment coefficients are : • Mean aerodynamic chord (MAC) = 27.58 in ches • Wing semispan = 115 .675 inches • Reference (planform) area of the semispan model = 2 , 973.6 in • Mo ment reference center (MRC): x = 132.59 inches, y = 46.875 inches , z = 17.795 inches 6 6 • B ased on MAC , Re = 3.27x10 for M = 0.20 (landing configuration) and Re = 4.24x10 for M = 0.26 ∞ ∞ (takeoff configuration) The schematic of the approximate pressure tap lo cation s , indicated by the red dotted lines, is shown in Fig. 6 .
Most of the pressure taps are in the streamwise arrays at 8 spanwise locations with 3 rows across the inboard flap span, 3 rows across the outboard flap span, and 2 rows across the aileron re gion. Additional ly, 6 spanwise arrays are on the upper wing surface with 1 row on the slat, 3 rows on the main wing , and 2 rows on the flap. The total number of pressure taps on the wing is approximately 6 80 . Othe r locations for pressure taps include the fuselage (~100), nacelle / pylon (~ 10 0), and horizontal stabilizer (~20) . In total, there are approximately 9 00 pressure taps on the configuration . There will also be approximately 100 unsteady pressure sensors installed on the model . The design of the slat and flap brackets is ongoing. Once the bracket number and location are decided, the pressure tap location s and number will be further adjusted to avoid any bracket interference.
Table 3 summarizes the slat and flap rigging s for the co nventional and the simpl e hinged flap high - lift configurations at the MAC (y = 46.875 inches) . For the conventional HL - CRM geometry , t he nominal slat and flap deflection s are 22 ° and 25 ° , respectively, for the take off configuration , and 30 ° and 37 ° , respe ctively, for the landing configuration . For the simple hinged flap HL - CRM geometry, the nominal slat and flap deflections are 30 ° and 50 ° , respectively, for the landing configuration. Th e nominal flap and slat deflections a r e generally the same for both inboard and outboard slats and flaps. The engine nacelle location is the boundary for the inboard and outboard slat s and the yehudi break is the boundary for the inboard and out board flap s . The gap and overhang definition is the 4 3 same as that reported by Lin and Dominik (see Fig. 7). Notice that one should use local wing chord as the reference chord ( C ) for the slat and the outboard flap , and use wing chord at the yehudi break as C for the ref ref inboard flap. Typically, t he range of adjustability for the slat gap, s lat overhang, and flap overhang is ~3% C . T he ref flap gap range of adjustability is ~ 1.5% C . The flap gap is fairly constant at 1.25% C and 0.9% C for the ref ref ref 2 4 landing and takeoff configurations, respectively, across the model span , as reported by Lacy & Sclafani. Ho wever, there is m ore variability for the flap overhang, slat gap, and slat overhang across the span .
Figure 6 . Schematic of pressure tap location s (as indicated by dotted red lines ) and cross section of the represent ative high lift conf igurations .
Table 3 . Slat and flap rigging for the HL - CRM configurations at MAC ( y = 46.875 inches ) .
Conventional Simpl e Hinged Flap (Baseline) Configuration Range Range Nominal Nominal ( Estimation) (Estimation) 27° to 33° 27° to 33° Landing 30° 30° (3° increment) (3° increment) Slat Deflection 19° to 25° 19° to 25° Takeoff 22° 22° (3° increment) (3° increment) Landing 1. 0 % C 0 to 3 % C 1. 0 % C 0 to 3 % C ref ref ref ref Slat Gap Takeoff 0 . 3 % C 0 to 3 % C 0 . 3 % C 0 to 3 % C ref ref ref ref Landing - 0 . 8 % C - 1 to 2% C - 0 . 8 % C - 1 to 2 % C ref ref ref ref Slat Overhang Takeoff 1. 5% C - 1 to 2 % C 1. 5% C - 1 to 2 % C ref ref ref ref 3 4 ° to 43 ° 4 0° to 6 0° (10° Landing 3 7 ° 5 0° Flap Deflection (3° increment) increment) Takeoff 25° 10° and 25° 25° 10° and 25° Landing 1 . 2 5% C 0 to 1.5 % C N/A N/A ref ref Flap Gap Takeoff 0. 9 % C 0 to 1.5 % C N/A N/A ref ref Landing 1. 2 % C 1 to 4 % C N/A N/A ref ref Flap Overhang Takeoff 3 . 3 % C 1 to 4 % C N/A N/A ref ref Note: C = local wing chord for t he slat and outboard flap, and C = wing chord at the yehudi break for the ref ref inboard flap.
4 3 Figure 7 . Gap and overhang definition for multi - element high - lift configuration.
Figure s 8 to 12 display some CAD images of the model . The cen ter spar layou t is shown in Fig. 8 . For illustration standardization purposes, the image of Fig. 8 is flipped from that of Fig. 6 such that the flow direction is from left to right. The spar will be hollow to allow for routing of instrumentation and AFC plum b ing . The model will have the capability for testing with and without the engine nacelle. There will be a filler piece to connect the inbo ard and outboard slat s when the engine nacelle is not installed, as shown in Fig.
9 . Figure s 10 and 11 show the image s of the tunnel f loor without any acoustic treatment for the aerodynamic testing and with the acoustic treatment for the aeroacoustic testing, respectively . Notice that the model will have a removable horizontal tail as a test option (see Fig. 10). The aeroacous tic testing requires a deeper floor due to the acoustic treatment, and thereby needs a specially made balance extension (see Fig. 11 ). The model will have a double - hinged mount at the model/balance interface to enable the model to be tilted to a horizonta l position (in both directions) for more efficient model change s and to enhance test productivity .
Figure 8 . HL - CRM modular features .
Figure 9 . HL - CRM with and without engine nacelle.
Figure 10 . HL - CRM with re g ular (non - acoustic) flooring.
Figure 11 . HL - CRM with acoustic flooring .
Project Schedule A schedule summary of the HL - CRM project is shown in Fig. 12 . The model design is ongoing, and its completion date is estimated to be March 31, 2017. There is a 12 - month period for model fabri cation. The w ind tunnel t esting i s scheduled to start around June 201 8 and end in March 2019 .
After the results for the AFC - enabled simpl e hinged flap high - lift test are analyzed, a decision will be made on whether to move forward with a la r ger scal e (~ 30%) model wind tunnel test and/or a flight test. In addition, if the effort to develop a low pneumatic power AFC system is successful, it may enable AFC applications on advance d aircraft designs that use simpl e hinged flap s (e.g., B lended W ing B ody ) and provide ( d ual - use ) AFC opportunities for high performance aircraft .
Figure 12 . Timeline for the development and testing of the HL - CRM .
Test Plan There will be ~32 weeks of tunnel occupation time at the 14x22. Aerodynamic test s will be conducted in t he first half (~14 weeks) with closed sidewalls, and aeroacoustic test s will be conducted in the second half (~14 weeks) with opened sidewalls. A 4 - week period is required for reconfiguring the wind tunnel from aerodynamic mode to acoustic mode. The aero dynamic test ing will include a rigging study that includes adjustment s to the flap /slat gap, overhang, and deflection setting s in order to refine the nominal conventional (baseline) high - lift configuration.
L ocalized AFC for L/D enhancement , AFC - enhanced simpl e hinged flap, and nominal conventional configuration with SCF and SGF devices will be tested in the aerodynamic portion. In addition to examining the effectiveness of various noise reduction devices, t he aeroacoustic test ing will also include acoust ic measurement s of the nominal conventional configuration and some selected AFC - enabled configurations. Table 4 illustrates the test plan for the HL - CRM test in the 14x22.
Table 4 . Test p lan for HL - CRM at the 14x22.
Model Estimated Test Description Measurement Techniques Configuration Duration Rigging study for B alance, surface pressures Takeoff and b aseline conventional (both steady and unsteady), 4 weeks landing HL - CRM tufts, model deflection using videogrammetry, Localized AFC for Takeoff 2 weeks stru cture vibration using L/D enhancement laser vibrometer , flow AFC - enhanced simpl e Aero dynamic Testing visualization using tufts hinged flap with Most ly landing 6 weeks (Closed sidewall) and IR cameras , and conventional slat optional PIV (schedule and Conventional HL - resource allow) CRM with SCF and Takeoff and 2 weeks SGF devices for slat landing gap noise reduction Wind tunnel test section reconfigured from aerodyna mic mode to aeroacoustic mode 4 wee ks (from closed sidewall s to opened sidewall s ) Baseline conventional Acoustic array, force Takeoff and HL - CRM (nominal balance, surface pressures 2 weeks landing configuration) (both steady and un steady), tufts, model deflection Slat gap noise using videogrammetry, reduction using SCF structure vibrati on using and SGF, well as Takeoff and Aeroacoustic Testing 10 Weeks laser vibrometer , flow optional flap edge landing (Opened Sidewall) visualization using tufts noise reduction and IR cameras , and devices optional PIV (schedule and Selected AFC - en abled resource allow) simpl e hinged flap Mostly landing 2 weeks configuration s Concluding Remarks This paper summarize s the current effort s to develop and design a 10% - scaled semispan HL - CRM for testing at the NASA LaRC 14x22 during FY 2018 . Both con ventional and AFC - enabled high - lift configurations are being designed and are intended to be open geometr ies . In addition to the current research effort, t he HL - CRM offer s numerous opportunities for high - lift R&D . Here are some potential uses for conside ration: 1. The HL - CRM provides a generic , open geometry high - lift model for ongoing testing in the 14x22.
2. The model can be used as a common test bed for advanced AFC actuators and noise reduction concept stud ies .
3. Since most of the wing components are modular, the model can be used to study wing geometry variations by replacing these modular components.
4. The HL - CRM can be used for c ollaboration opportunities with industry, academia, and other government agencies through cooperative agreements such as Space Act A greement s (SAA), NASA Research Announcement s (NRA), Small Business Innovation Research (SBIR) projects , or Interagency Agreement s (IA).
5. B ecause it i s an “open” high - lift geometry , t he HL - CRM design is an excellent candidate for CFD code validation activiti es such as the CFD High - Lift Prediction Workshop .
6. The model and geometry can be used for Federal Aviation Administration ( FAA ) related R&D, such as icing stud ies .
Acknowledgments This R&D effort involved many people who provided valuable contributions th rough their detailed work and/or expert guidance. The contributors have come from within NASA and Boeing. Boeing developed the CRM high - lift geometry through a combination of internal R & D and NASA task orders (2013 - present). NASA contributions have been sponsor ed by the Fixed Wing (FW) P roject from FY 2013 - 2014 and Advanced Air Transport Technology (AATT) Project from FY 2015 - present. The authors would like to thank all those involved for their great support during the course of this R&D effort . Some k ey personnel are as follows : NASA Langley Research Center : Craig H unter and Melissa Rivers ( CFD ); Dave Castle, Mark Cagl e, Sandy Webb, and Reggie Kidd (model design) ; Dave Lockard, Travis Turner , Meela n Choudhari, Mehdi Khorrami, Florence Hutcheson, and C raig Streett ( aeroacoustic ) ; Norma Farr (model geometry); Dan Neuhart ( model instrumentation ); Bill Jones and Chris Rumsey (CFD workshop) ; Doug Nark and Hamilton Fernandez (aeroacoustic managers); S usan Wilz a nd Scott Anders (AATT managers); Rich Wahls ( former FW manager) .
Boeing Research & Technology : Ed Whalen and Abe Gissen ( AFC study manager s ) .
Boeing Engineering Operations & Technology : Arvin Shmilovich, Yoram Yadlin , Eric Dick e y , and Pichuraman Sundaram (CFD); Peter Hartwich (system study manager) ; and Abd i Khodadoust ( Boeing/ NASA contract manager) .
Boeing Commercial Airplanes : Doug L acy , P aul Vij gen , and To ny Sclafani (high - lift aerodynamics).
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