# AE2610 Introduction to Experimental Methods in Aerospace AERODYNAMIC FORCES ON A WING IN A SUBSONIC WIND TUNNEL

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1 AE2610 Introduction to Experimental Methods in Aerospace AERODYNAMIC FORCES ON A WING IN A SUBSONIC WIND TUNNEL Objectives The primary objective of this experiment is to familiarize the student with measurement of aerodynamic forces in a wind tunnel. A six-component sting balance will be used to measure forces and moments on a rectangular (two-dimensional) wing with an adjustable plain flap. In addition, the student will become acquainted with the operation of a subsonic wind tunnel, including use of a Pitot-static probe and pressure transducer to measure wind speed. Background a) Aerodynamics of a Wing i) Forces and Moments In general, a body moving through a fluid is subject to three force and three moment components. The primary force components on an aerodynamic body, and in particular a wing, are lift and drag. Lift is the force component that is perpendicular to the oncoming flow direction, and drag is the component parallel to the flow (see Figure 1). Lift is generally considered positive when in a direction upward (opposing gravity), while drag is normally positive in the flow direction (i.e., tending to slow down the body). The third component is the called the side force. For objects symmetric about the side force axis and moving straight into the flow, there should be no side force. Side forces on a wing usually come about when an aircraft is turning or not flying into the wind. The three moments are: the pitching, roll, and yaw moments (see Figure 1). The pitching moment is around an axis parallel to the direction of the side force, and would act to change the pitch or angle of attack ( ) of the lift body. The roll moment is around an axis in the drag direction, while the yaw moment acts around the lift axis. For most wings flying into the wind, the pitching moment is the dominant component. It is customary to define the pitching moment about the aerodynamic center; the point at which the pitching moment does not vary with lift coefficient, i.e., angle of attack. The magnitudes of the aerodynamic forces and moments depend primarily on the shape of the body, the speed and orientation of the body with respect to the fluid, and certain properties of the fluid. For a rectangular wing, lift primarily depends on the curvature (camber 1 ) of its 1 The camber line of a wing s cross-section connects the midway points between the upper and lower surfaces. Copyright , 2002, 2009, 2015, 2018, 1 by H. McMahon, J. Jagoda, N. Komerath, and J. Seitzman. All rights reserved.

2 cross-section, the angle of attack, the flow speed, and the density of air. The shape of the wing s cross-section is known as an airfoil. The thin airfoil inviscid flow model 2 (also called thin airfoil theory ) predicts that lift increases with increasing positive (concave downward) camber or with rearward movement of the location of the maximum camber. In a real viscous airflow, this is also true at moderate angles of attack. Thus, the maximum value of the lift coefficient attainable with any airfoil increases with increasing camber. This is of great benefit, since the higher the maximum lift coefficient the lower the stalling speed. Also, the effective camber of an airfoil can be varied by deflecting a segment of the airfoil such as a flap or aileron. It may be shown that, for a given amount of deflection, the change in lift coefficient is larger if the deflection occurs at the trailing edge rather than at the leading edge. Hence, flaps and ailerons are fitted at the trailing edge of an airfoil or wing. Additionally, thin airfoil theory predicts that the aerodynamic center of an airfoil is located one-quarter of a chord length behind its leading edge. The rectangular wing used in this experiment (see Table 1 for dimensions) is fitted with a plain trailing edge flap. Aerodynamics forces on the airfoil will be measured as a function of wing angle of attack and flap deflection angle. The lift results will show if increased flap deflection has the expected effect, while the drag results will show whether an increase in lift at a fixed angle of attack by means of a flap deflection is accompanied by any corresponding drag change. If there is a drag increase this is not necessarily a disadvantage. For example, during landing the higher engine power required to compensate for the drag increase will minimize engine acceleration time in the event of a missed approach. ii) Aerodynamics Coefficients Aerodynamics forces on a wing are typically normalized and expressed as aerodynamic force coefficients. For example, the lift coefficient (CL), drag coefficient (CD) and pitching moment coefficient (CM) are defined as C L C D C M L (1) q S D (2) q S M (3) q Sc 2 See for example, Anderson's Fundamentals of Aerodynamics. 2

3 where L is the lift force, D is the drag force, M is the pitching moment, S is the plan form area of the wing, c is the chord length (see Figure 1), and q is the dynamic pressure. For our rectangular wing, S = c b, where b is the span (tip-to-tip length) of the wing. Thin airfoil theory predicts that CL for a symmetric airfoil, i.e., where the chord and camber lines are the same (also known as a zero camber airfoil) that also has infinite span is given by C 2 (4) L This expression is valid only for small and moderate angles of attack. At sufficiently large angles of attack, the flow over the wing no longer follows the surface, and the wing s lift coefficient begins to drop, i.e., the wing stalls. The dynamic pressure q is given by q (5) v where is the density of the approaching flow and v is its speed (relative to the body/wing). This is called the dynamic pressure because according to Bernoulli's equation (which is valid for a low speed, constant density flow) it represents the difference between the stagnation (or total) pressure and static pressure of a flow, i.e., p o p (6) v The static pressure is the pressure one would measure if moving with the flow (or without requiring a change in the flow velocity), while stagnation pressure is the static pressure that the flow would achieve if it was slowed (in an ideal manner) to zero velocity. For an ideal gas, the density is given by p (7) RT Here p is the (absolute) pressure, T is the (absolute) temperature, and R is the gas constant for the specific gas. b) Sensors and Transducers i) Sting Balance The force-measuring device to be used in these experiments is a sting balance mounted on a model positioning system (MPS) that controls both pitch and yaw angles. Wind tunnel force balances can be categorized into two basic types: external and internal. The force transducers for external balances are located outside the wind tunnel test 3

4 section, and the aerodynamic forces must be mechanically transmitted from the test body in the wind tunnel to the balance outside. Internal balances have their force measuring unit within the wind tunnel, and thus provide a more direct measurement of the aerodynamic forces. The Aerolabs TM internal balance (Figure 2) employed in our wind tunnel uses strain gauges mounted within a rod cut from a single piece of precipitation-hardened stainless steel with a protective sheath to protect the fragile foil strain gauges. The strain gauges are connected to electronics (including wheatstone bridges), which are in turn connected to a computer data acquisition system located in the wind tunnel control room. The Aerolab system will acquire 5000 samples every 100 ms and report the average. By setting the dwell time of the MPS, you can set the number of measurements that will be recorded at each angle of attack (e.g., a dwell time of 1 second would result in 10 measurements). Two procedures are required to convert the voltage outputs from the strain gauge system into accurate measurements of the aerodynamics forces (and moments). First, the strain gauges must be calibrated. This is accomplished using an attachment of known mass that can be connected to the balance. Second, any forces transmitted to the balance that are not produced by the wing must be removed. In this experiment, there is a support structure that connects the balance to the wing (see Figure 3). This support structure can produce its own aerodynamics forces. In addition, even with the wind tunnel off, the weight of the structure and wing will load the balance. Removing these wind-off (or tare) loads is known as taring the system. If the data are to be meaningful, each test run must be carried out at a (nearly) constant value of wind tunnel dynamic pressure. Ideally, the magnitude of the aerodynamics force coefficients would be independent of the magnitude of the dynamic pressure. In reality, this assumption might not be completely accurate, so the dynamic pressure should be held constant during a particular data-taking run so that the coefficients will not be influenced by changes in freestream conditions. The speed of the fan that drives the wind tunnel is nominally held constant by the fan s motor control system. This should produce a nearly constant wind tunnel dynamic pressure. (If necessary, the dynamic pressure can be manually adjusted with the tunnel speed control.) ii) Pitot-Static Probe - In addition, a second data acquisition system will simultaneously acquire and store the dynamic pressure during the experiment. The dynamic pressure is measured directly by means of a Pitot-static probe mounted in the freestream. The probe has one hole facing directly into the wind tunnel flow and another located so the air flows along the surface of the hole (see Figure 4). The pressure in the first hole is the stagnation pressure (po) and the second experiences the static pressure (p). The two holes are connected via long 4

5 lengths of tubing to the two sides of a capacitance-type differential pressure transducer (Baratron). This transducer interprets the displacement of a diaphragm due to a change in pressure difference across the diaphragm as a change in capacitance in an electronic circuit, which is output as a DC voltage change. This DC voltage is displayed on a digital voltmeter, which can be monitored by the wind tunnel operator (as noted above, it is also connected to the data acquisition system). The output voltage of the Baratron can be converted to a pressure using the sensitivity of the transducer (e.g., mmhg/volt). c) The Wind Tunnel A wind tunnel is a duct or pipe through which air is drawn or blown. The Wright brothers designed and built a wind tunnel in The basic principle upon which the wind tunnel is based is that the forces on an airplane moving through air at a particular speed are the same as the forces on a fixed airplane with air moving past it at the same speed. Of course, the model in the wind tunnel is usually smaller than (but geometrically similar to) the full size device, so that it is necessary to know and apply the scaling laws in order to interpret the wind tunnel data in terms of a full scale vehicle. The wind tunnel used in these experiments is of the open-return type (Figure 5). Air is drawn from the room into a large settling chamber (1) fitted with a honeycomb and several screens. The honeycomb is there to remove swirl imparted to the air by the fan. The screens break down large eddies in the flow and smooth the flow before it enters the test section. Following the settling chamber, the air accelerates through a contraction cone (2) where the area reduces (continuity requires that the velocity increase). The test (working) section (3) is of constant area (42" 40"). The test section is fitted with one movable side wall so that small adjustments may be made to the area in order to account for boundary layer growth, thus keeping the streamwise velocity and static pressure distributions constant. The air exhausts into the room and recirculates. The maximum velocity of this wind tunnel is ~50 mph, and the turbulent fluctuations in the freestream are typically less than 0.5% of the freestream velocity. Thus, it is termed a low turbulence wind tunnel. 5

8 use the wind tunnel controller to set the wind tunnel to the desired speed. DO NOT hit start on the Labview program yet! 12. Now you will record the aerodynamic forces with the tunnel on by following the same procedure outlined in step 10 - HOWEVER this time you will need to synchronize the start and end points of the acquisition on the Aerolab and Labview software by beginning both at the same time (this will require two people one working each computer). 13. Have the TA help with removing the test model from the MPS. Then attach the wing after setting the flap angle to zero. Then, have the TA help reinstall the model to the MPS. 14. Repeat steps Have the TA help with removing the test model from the MPS. Then set the flap angle to your chosen angle. Then have the TA help reinstall the model to the MPS. 16. Make sure the MPS is set to zero pitch (and zero yaw), and make sure the system is retared. 17. Repeat step 12 (i.e., a sweep for the chosen wind tunnel speed) at your new flap angle. 18. After the tests are concluded, the appropriate data files will be transferred by the TA s to the lab website for later data reduction (as noted below). DATA TO BE TAKEN 1. You must record the room temperature and barometric pressure in order to set the wind tunnel speed. 2. Tared forces and moments of the wing support structure at 10 angles of attack. 3. Tared forces and moments of wing and support structure at the same 10 angles of attack. DATA REDUCTION 1. Remove the support structure loads from the loads acquired with the wing + support structure to find the wing-only normal force, axial force and pitching moment. 2. Convert the measured forces in the balance reference frame to the wind frame (i.e., convert to lift and drag). 3. Convert the recorded pitch moments about the electrical center of the sting balance to the (nominal) aerodynamic center of the wing. The electrical center of the sting balance is located as shown in Figure Express all forces and moments in coefficient form. The chord to be used is the chord of the wing from the wing leading edge to the flap trailing edge. 8

9 RESULTS NEEDED FOR REPORT 1. Wing geometry - chord of wing (including flap), span, chord of flap, planform, thickness ratio. 2. Plot a figure showing lift coefficient (ordinate) versus angle of attack (CL vs. ) for the wing at the two flap deflection angles. 3. Plot a figure showing drag coefficient (ordinate) versus angle of attack (CD vs. ) for the wing at the two flap deflection angles. 4. Plot a figure with drag coefficient as abscissa and lift coefficient as ordinate (CL vs. CD also known as a drag polar) for the two flap deflections. 5. Plot a figure with moment coefficient (ordinate) versus angle of attack (CM vs. ) for the wing at the two flap deflection angles. Table 1. Dimension of the rectangular planform wing at zero flap angle. Dimension Value (inches) Span Chord (including flap) 3.00 Flap Chord 0.82 Maximum Thickness

10 Mean Camber Line Lift Yaw v Pitch Roll Drag c Chord Line Figure 1. Some of the force and moment components on an airfoil. The angle of attack ( ) is shown between the chord line and the flow direction. Figure 2. Schematic of Aerolab internal strain gauge sting balance (shown without protective sheath) mounted on model positioning system (MPS). 10

11 Figure 3. Schematic of wing and support structure connected to sting balance (left bottom view; right top view). p o p v Figure 4. Schematic of Pitot-static probe in a wind tunnel. 11

12 Figure 5. Georgia Tech low turbulence wind tunnel. 12

13 2.0 Wind AF 9.9 Electrical Center Figure 6. Electronic center of sting balance relative to wing leading edge. 13

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