Pressure Management on a Supercritical Airfoil

Pressure Management on a hypercritical control surface in tran transonic prevail Abstract-At transonic runs an open will throw f emit accelerate onwards from the leading march on to sonic stimulates and originate a seismic disturbancewave oer the surface of its body. One particularor that determines the impactwave location is the prevail speed. However, the shape of an control surface too has an influence. The experiment conducted comp ard Mach flow over a critical airfoil (planate upper surface) and a naca0012 aerofoil (symmetrical).Despite discrepancies, the experiment sustain the flowing performance of a critical aerofoil being hypernym to a accomplished aerofoil. A comparison of the graphical distributions demonstrates the more than(prenominal) until now wedge distribution on a critical aerofoil and a longer delay in shockwave shaping. All of which, reflects the theory. Table of Contents Introduction3 Apparatus3 innovation Wind Tunnel with Transon ic Test Section3 Mercury Manometer4 Procedure4 theory and Equations5 Results6 Discussion10 Theory of Transonic Flight10 Relating the Theory to the sample11Effectiveness of Supercritical aerofoils 11 Limitations and Improvements12 Appendix13 References14 Introduction For whatever prey break downling through a fluid such as air, a hug distribution over all of its surface exists which helps generate the obligatory lift. Lift is an aerodynamic force which is perpendicular to the direction of the aerofoil. Transonic speeds force in the formation of shockwaves over the top surface of the aerofoil. This is due to deepen flow over the surface region. We say this region is approximately amidst 0. 8-0. . Since the flow must accelerate and then will lose speeding following the shockwave the aerofoil will have a subsonic and sonic region. For the majority of commercial airlines this is non a desired region to cruise at given the instantaneous pressure distribution which passengers would otherwise scram. Particularly, the formation of shock shoot forth boundary form separation. Supercritical aerofoils ar more efficient knowing for higher Mach speeds and drag in reduction. They atomic number 18 distinct from established aerofoils by their flattened upper surface and asymmetrical design.The main payoff of this type of aerofoil is the development of shockwaves moreover away then tralatitious aerofoils and thus greatly reducing the shock induced boundary layer separation. In order to truly meet the effectiveness of a critical aerofoil an experiment gathering supercritical aerofoil performance and huffy data of a naca0012 aerofoil will be extensively analysed and comp atomic number 18d. quest the calculation and procedureit will be assessed whether a supercritical aerofoil is more effective. ApparatusA wind tunnel with a transonic hear section was employ in this experiment to study transonic flow around an aerofoil. The test section consists of lin ers which, after the initial contraction, are nominally parallel by from a slight divergence to compensate for growth of the boundary layers on the wall. In order to reduce interference and blockage at transonic speeds, the top and bottom liners are ventilated by longitudinal slots approve by plenum chambers. The working section has a height and width of 178mm and 89mm respectively. The doldrums pressure, p0? in the tunnel is close to atmospherical pressure, and therefore it keep be taken to be equal to the posittling-chamber pressure as the errors are scarcely small. To minimise the disturbance due to the model itself, the reference doldrums pressure, p? , is taken from a pressure tapping in the floor of the working-section, hygienic upstream of the model. The nominal free-stream Mach recite, M? , in the tunnel can be mensurable from the ratio p? /p0?. The Mach number in the tunnel can be controlled by varying the pressure of the injected air, pj. The maximum Mach number t hat the tunnel can achieve is about 0. 8 Mercury Manometer A multi-tube manometer with quicksilver was apply to measure the pressure at stagnation, the aerofoil tappings and atmosphere. The manometer is equipped with a lockup mechanism which allows the mercury levels to be frozen so that readings can be taken once the flow has been stopped. Also, the angle of the manometer can be adjusted. For this experiment, it was set to 45 degrees (Motellebi, F. ,2012). Procedure Before conducting the experiment, the barometric pressure, Pat, was recorded, in inches of mercury and the atmospheric temperature, in degrees Celsius, was also recorded.For a range of revalues of Pj from 10 cx lb/in2, in intervals of 20lb/in2, Pj was then recorded along with the manometer readings corresponding to stagnation pressure (I0? ), the reference static pressure (I? ), airfoil pressure tappings (In, n=1 to 8 and 3a) and the atmospheric pressure (Iat), all in inches of mercury (Motellebi, F. ,2012). Result s- Raw data in appendix x/c enroll 1b Cp against x/c at M= 0. 85 see 1a -Cp against x/c at M=0. 85 The data-based data was reborn to absolute pressure values using Equation x ( units are inches of mercury).For a given value of the pressure injector (Pinjector) we can find the value of the Mach number using Equation y. Also Equation Z calculates Cp( or pressure coefficents) which reflect the measurements of the surface of the aerofoil. These results are displayed in figure x. This was make for both the supercritical aerofoil and the NACCA 0012 aerofoil. What follows is a comparison and analysis of the data. ( image 2b Cp against x/c at Mach speed 0. 8 Figure 2a -Cp against x/c at Mach speed 0. 81 x/c x/c Figure 3b- -Cp against x/c at Mach speed 0. 72 Figure 3a Cp against x/c at Mach speed 0. 3 Figure 4b Cp against x/c at Mach speed 0. 61 Figure 4a Cp against x/c at Mach speed 0. 61 Figure 5a- -Cp against x/c at Mach speed 0. 45 Figure 5b- -Cp against x/c at Mach speed 0. 44 Note that for both supercritical and naca0012 aerofoils the supercritical cases ( where M is equal to 0. 77, 0. 83 and 0. 840) the approximate value of x/c % where the shock occurs over the aerofoil is shown in red line. For the point below where Cp and the Cpcritical and accordingly the drop in Cp is greatest gives the location of where the shockwave occurs on the surface of the aerofoil. Cp and Cp* vs M? naca0012 aerofoil) Cp and Cp* vs M? (supercritical aerofoil) It is worth noning that for both the supercritical and Naca0012 aerofoil the results are somewhat similar. That is the critical Mach amount for both are around 0. 72. therefore the borderline Mach number for a local shockwaves on both the supercritical and conventional aerofoil can be assumed to be the same. It is worth noning that Mach number 0. 41 for the supercritical aerofoil does not produce a shockwave, whereas the Naca0012 aerofoil does. Mach number Supercritical Aerofoil Approx position of shock naca0012 Appro x position of shock 0. 5 - - 0. 61 - - 0. 72-0. 73 - 0. 25x/c% 0. 85-0. 86 0. 70x/c% 0. 40x/c% Basic transonic theory An aerofoil or any purpose for that matter travelling through a medium (air) at low Mach numbers ( typically between 0. 30-0. 40) has flow is subsonic and can be considered incompressible. This message that any change in pressure or concentration is significant. The speed of sound (a) is dependent on the altitude of the aerofoil/objective and the Mach number M is the ratio of velocity M=va , a=? RT ?is a specific heat ratio, T is thel absolute temperature and R is the gas constant.The conspiracy of these 2 equations above leads to M=v? RT Sound is essentially a series of consecutive weak pressure waves emitted from a given spring. These waves travel at the local speed of sound. If we assume the aerofoil is travelling towards the source, the source can notice the disturbances beforehand giving enough time for flow to adjust around the object. When the source beg ins to approach near the speed of sound, pressure waves move closer together in front of the object, therefore incompetent information from the source/disturbance is propagated upstream and the flow will not be able to react in time.The pressure waves merge together to produce a shockwave in front of the object. The flow encountering the shockwave will experience changes in temperature, static pressure and gas density as well as a lower Mach number. The transonic region is special because although flight speed is below sonic speed as the information is propagated upstream on the surface of the aerofoil the flow accelerated to the speed of sound. and so forming a shockwave over the aerofoil. The position of this shockwave depends on the initial entry speed to the aerofoil.Therefore what we have in the transonic region is an aerofoil which has sonic speeds azoic upstream and subsonic speed towards the end of the aerofoil or downstream. This means it is complicated to accurately a nalyse transonic flow over an aerofoil as a disparate set of equations must be employ on the leading edge, upper surface and trailing edge. The critical upstream Mach number is the minimal value of a given Mach number for which a shockwave will be produced on the surface of an aerofoil. In other words, supersonic flow.Below this threshold a shockwave will not appear. Drag or the aerodynamic force in the transonic region again depends on the speed of the object travelling. At subsonic speeds the main component of drag are Skin friction, pressure drag and lift induced drag. At sonic speeds (approaching or exceeding) there is the addition of wave drag. The drag increases dramatically, and as a result a higher thrust is needed to sustain acceleration. Also, at this point the shockwave will interact with the boundary layer, thus causing it to bring out upstream of the shock.Figure 6Demonstration of transonic flight-(Scott, J. , 2000) The aerofoils The two aerofoils Naca0012 and Supe rcritical aerofoil are varied in design and purpose. The Naca0012 is a basic symmetrical aerofoil utilise primarily for rudder and elevator movements. Aerodynamic performance is not taken into term and is thus reflected by the simple aerodynamic design. It is worth noting that there are better aerofoils. The Supercritical aerofoil is a performance aerofoil designed for higher Mach speeds and drag reduction.They are distinct from conventional aerofoils by their flattened upper surface and asymmetrical design. The main advantage of this type of aerofoil is the development of shockwaves further away then traditional aerofoils and thus greatly reducing the shock induced boundary layer separation. Relating the Theory to the Experiment The critical Mach number for both the supercritical aerofoil and NACA0012 aerofoil was open to be in the region of 0,72. There is a difference to the close 10th but for all intents and purposes we can assume they are the same.This indicates that the mi nimum Mach number for a shockwave to be produced on the surface of the aerofoils is equal and not influenced via the shape. The pressure distributions of the supercritical aerofoil ( especially at Higher Mach) in comparison to the Naca0012 are more evenly distributed. The experiment affirms the theory that the supercritical aerofoil in comparison ro a conventional aerofoil generates more lift due to an even distribution of pressure over the upper surface. (http//en. wikipedia. org/wiki/Supercritical_airfoil) Effectiveness of Supercritical aerofoils.At a Mach number of 0. 45 both aerofoils do not display a shockwave. This is evident from the fact the Cp and Cp* graphs do not intersect at all. We already know this because the critical Mach number is 0. 72 for both. This indicates that either a shockwave was not produced (unlikely), or that the shockwave was produced beyond the trailing edge This means we cannot assess the effectiveness of the supercritical aerofoil at Mach speeds 0,4 5 and 0. 61. The supercritical Mach numbers show varying results. When the experiment took place at Mach ) 0. 72-0. 3 ( the critical Mach number) the supercritical aerofoil did not produce a shockwave ( Cp and Cp* do not intersect) whereas the naca0012 aerofoil did. The lack of a shockwave formation indicates either the critical Mach number for the supercritical aerofoil is higher then the conventional aerofoil experimental accuracy is lacking. At the supercritical mach numbers ( 0. 81-0. 86) in both the naca0012 aerofoil and the supercritical aerofoil Cp and Cp* intersect. The large drop in pressure coefficient is evidence of the formation of a shockwave.However, the pressure drop in the supercritical aerofoil is occurring at a pressure tapping further downstream. This confirms the theory that a shockwave is produced further downstream in a supercritical aerofoil This seems to confirm the theory that a supercritical aerofoils design does allow for development of shockwaves further away then traditional aerofoils and thus greater reduction in the shock induced boundary layer separation. In regards to the amount of drag (aerodynamic force) acting on the aerofoils it is worth noting that the pressure distribution at 0. 5 Mach for the supercritical aerofoil is more evenly distributed and flatter then the naca0012 aerofoil. There is no indication of a large instantaneous increase in drag taking over. This would therefore confirm the theory that a supercritical aerofoil is effective in greatly reducing the shock induced boundary layer separation. Notes for limitations The experiment is a success since results obtained confirm the capabilities of supercritical aerofoils and their advantages over conventional aerofoils. However, there are a few discrepancies which regarding experimental error and the different aerofoils.First of all the mach numbers tested at 0. 72 and 0. 73 created an inaccurate experiment. Normally, this would not be a problem. However, since the c ritical mach numbers for both aerofoils were in the vicinity of 0. 72 it was evaluate this was the minimum threshold for a shockwave to be produced over the aerofoil. A shockwave was not produced for the supercritical aerofoil despite the critical mach number value. Therefore, we can decide that at this speed there are too many inaccuracies to understand what is really going on.We also did not really see a difference in performance at subsonic flow. Granted, the supercritical aerofoil was primarily designed for supercritical mach speed. No useful information was obtained from here. The fact the pressure tappings have different coordinates means that each aerofoil is screening the pressure distribution at a different set of coordinates. This of course, is not as accurate if the aerofoils had the same pressure tappings. For instance, the naca0012 has a pressure tapping at 6. 5% of the aerofoil section and the last ends 75% the rest is unaccounted for.Since the supercritical aerofoi l has different pressure tappings it means both aerofoils have different areas which are unaccounted for. This means it is not certain whether or not the graphs are a reliable source of information, yet alone to compare. A digital meter should also be connected that displays the pressure in the two tappings so the aerofoil can be appropriately adjusted to bring it to zero incidence. This digital meter can also be employ to display the value of the mercury levels for other pressure tappings, reducing any human errors.In order to increase the accuracy of the pressure distribution over the aerofoil surface, more pressure tappings can be made on the aerofoil. These will improve the pressure coefficient graphs by allowing more points to be plotted, in turn, yielding better information for the position of the shockwave in the supercritical cases and also the critical Mach number for a shock to occur. References 1) http//www. southampton. ac. uk/jps7/Aircraft%20Design%20Resources/aerodyna mics/supercritical%20aerofoils. pdf 2) http//www. nasa. gov/centers/dryden/pdf/89232main_TF-2004-13-DFRC. pdf 3)