8. PERFORMING ORGANIZATION

Transonic steadyand unsteady-pressure tests have been conducted on a large elastic wing. The wing has a supercritical airfoil, a fullspan aspect ratio of 10.3, a leading-edge sweepback angle of 28.8 , and two inboard and one outboard trailing-edge control surfaces. Only the outboard control surface was de ected statically and dynamically to generate steady and unsteady ow over the wing. This report presents the unsteady-surface-pressure and dynamic-de ection measurements of this elastic wing, in tabulated form, to permit correlations of the experimental data with theoretical predictions. Introduction At the NASA Langley Research Center, progress continues on a program to obtain measured unsteady pressures on several di erent wing con gurations (refs. 1{3). The goal of this program is to generate an extensive data base of measured unsteady pressures for use in evaluating the accuracy of theoretical computational transonic aerodynamic programs. Initially, all the wing models that were tested were made as rigid as possible to minimize wing structural deformations and thereby maintain simple basic comparisons with the transonic aerodynamic programs. Recently, a exible wing con guration was tested as part of this pressure measurement program. The exible wing construction is similar to that of actual aircraft wings and should provide more realistic measured data for comparison with the results from the advanced transonic aerodynamic programs including the e ects of aeroelastic deformations in the computational process. This elastic wing con guration, known as the Drones for Aerodynamic and Structural Testing Aeroelastic Research Wing-2 (DAST ARW-2, ref. 4), has a full-span aspect ratio of 10.3 (excluding the area of the wing trailing-edge extension), a leadingedge sweepback angle of 28.8 , and a supercritical airfoil. The wing has three hydraulically actuated trailing-edge control surfaces and is instrumented with unsteady-pressure gages, making it extremely useful to the present unsteady-pressure-measurement program. The two inboard control surfaces were held xed while the outboard control surface was oscillated to create the unsteady pressures. This report is one of a series of reports documenting the data acquired on the DAST ARW-2 (refs. 5{9). The purpose of this report is to document, for future use, the measured unsteady-pressure and wingde ection data results from an elastic wing con guration tested in the Langley Transonic Dynamics Tunnel (TDT). All pressure results are tabulated and presented in pressure-coe cient form. Symbols ACC MAG magnitude of wing accelerometer signal, G units AMPL amplitude of oscillations, deg b semichord at y = 0, in. (22.12 in.) CP pressure coe cient, (p P )=q CPSTAR critical pressure coe cient DELTA CP lifting-surface pressure coe cient, lower surface CP Upper surface CP f frequency, Hz G = z=g g gravity constant, 386.088 in/sec H stagnation pressure, psf K reduced frequency, b! V MACH free-stream Mach number P free-stream static pressure, psf p local static pressure at any point on wing surface, psf q free-stream dynamic pressure, psf (Q in computer-generated tables) RN Reynolds number based on average chord of 24.812 in. V free-stream velocity, in/sec X streamwise distance measured from wing local leading edge, in. X=c fraction of local-chord location (X/C in computer-generated tables and gures) x streamwise coordinate, in. y spanwise coordinate, in. z wing vertical de ection amplitude, in. z vertical acceleration, in/sec wing angle of attack, positive for leading edge up, deg (ALPHA in computergenerated tables) ratio of speci c heat at constant pressure to speci c heat at constant volume (GAMMA in computer-generated tables) control-surface angle about hinge line, positive for trailing edge down, deg (DELTA in computer-generated tables) fraction of wing semispan (ETA in computer-generated tables) ! oscillation frequency, rad/sec Wind Tunnel Model General An elastic semispan wing model is described herein. This model consisted of the right wing panel from the Drones for Aerodynamic and Structural Testing Aeroelastic Research Wing-2 (DAST ARW-2) drone ight vehicle and a rigid half-body fuselage. Both the fuselage and the wing were mounted on a remotely controlled turntable mechanism located on the tunnel sidewall. The turntable was used to adjust the model angle of attack. A photograph, looking upstream, of the complete model mounted in the tunnel is shown in gure 1. The location of the sidewall turntable and its relationship to the wing and fuselage are shown in gure 2. For all the tests contained in this report, no boundarylayer trips were used; the boundary-layer transition on the wing was left free. Fuselage Geometry and Construction The geometric shape of the fuselage is shown in gure 2. Fuselage coordinates and further details about the structure are given in reference 8. The rigid half-body fuselage was used primarily to place the wing outside the wind tunnel wall boundary layer. The fuselage had a semicircular cross section. The nose and tail fuselage sections were made shorter than the actual ight fuselage. However, the center section of the fuselage was made very similar to the ight fuselage in both diameter and wing location to provide ow around the inboard section of the wing similar to that expected to occur on the ight vehicle. This fuselage shape represents that of a typical transport aircraft. Wing Geometry, Construction, and Structural Properties The elastic wing had a full-span aspect ratio of 10.3 (excluding the area of the wing trailingedge extension) with a leading-edge sweepback angle of 28.8 . The planform geometry of the wing is presented in gure 3. The wing was equipped with three hydraulically actuated control surfaces, two inboard and one outboard. Their locations are also shown in gure 3. Only the outboard surface was de ected statically and dynamically during the pressure-measurement tests while both of the inboard surfaces were held xed at 0 in relation to the wing surface. The outboard surface hinge line was located at 77 percent of the local chord. The wing contour was the desired shape for a loaded wing associated with straight and level ight of a vehicle at a cruise Mach number of 0.8 and at an altitude of 46800 ft with a lift coe cient of 0.53. However, an elastic wing will deform to a di erent shape, known as the jig shape, if all aerodynamic loads and vehicle weight loads are removed. The present wing con guration was fabricated to a set of calculated jig shape coordinates referred to as the design airfoil coordinates. Design coordinates and the measured coordinates from the actual wing cantilevered at the root chord are available from table 4 of reference 8. A detailed description of the wing construction, including how the calculated jig shape was determined, is found in reference 8. Also, reference 8 contains a detailed description of the structural properties of this elastic wing along with a structural nite-element model. Instrumentation The locations of the wing instrumentation are shown in gure 3. The primary instrumentation consisted of 182 pressure transducers and 10 accelerometers. In addition, strain gage bridges were located near the wing root to measure bending moments. A di erential pressure gage was mounted in each supply line to the hydraulic actuator of each control surface to measure hinge moments. Small potentiometers 2 were used to measure the control surface angular displacement. The model angle of attack was measured by a servo accelerometer that was mounted near the wing root. Both steady and unsteady surface pressures were obtained with di erential pressure transducers referenced to the static pressure of the tunnel. Streamwise rows of upperand lower-surface ori ces were located at six span stations. The wing location of these ori ces is given in table 1. Steady pressures were measured at all six span stations. Unsteady pressures were measured on only the three outermost span stations. Surface ori ces were connected to pressure transducers by matched tubes (ref. 10) having an inner diameter of 0.040 in. and a length of 18 in. To determine the wind-on tube transfer functions that are needed to correct the unsteadypressure data from these matched-tube transducers, simultaneous measurements were also obtained from a row of in situ transducers (see g. 3) mounted on the wing upper surface parallel to the fth row of surface ori ces. Based upon the manufacturer's speci cations, the unsteady-pressure transducers used are accurate to within 0.038 psi. The 10 accelerometers were used to determine the wing dynamic de ections. The accelerometer locations are shown in gure 3 and presented in table 2. The accelerometers were mounted in the wing approximately halfway between the upper and lower surfaces. Wind Tunnel The tests described in this report were conducted in the Langley Transonic Dynamics Tunnel (TDT). The TDT is a closed-circuit, continuousow tunnel that has a 16-ft square test section with cropped corners and with slots in all four walls. Mach number and dynamic pressure can be varied simultaneously or independently, with either air or a heavy gas used as a test medium. A heavy gas was used as the medium for all the tests contained in this report. Data Acquisition and Reduction All data from the model instrumentation were acquired with the TDT real-time data-acquisition system (ref. 11). The pressure measurements were acquired with an electronically scanned pressure (ESP) system (ref. 12). The ESP system is a sequential, digital pressure sampling equivalent to a mechanical scanivalve. The pressure data were digitized in real time at 250 samples per second and written on magnetic tape for later analysis. Unsteady pressures were measured for 90 ESP pressure transducers and 7 in situ pressure transducers. The accelerometer and control surface position data were acquired simultaneously, digitized in real time at 1000 sample