Performance Characteristics of Flush Angle-of- Attack Measurement System Integrated on a Pitot Tube

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1 Engineering Applications of Computational Fluid Mechanics ISSN: (Print) X (Online) Journal homepage: Performance Characteristics of Flush Angle-of- Attack Measurement System Integrated on a Pitot Tube J. Masud To cite this article: J. Masud (2010) Performance Characteristics of Flush Angle-of-Attack Measurement System Integrated on a Pitot Tube, Engineering Applications of Computational Fluid Mechanics, 4:4, , DOI: / To link to this article: Copyright 2010 Taylor and Francis Group LLC Published online: 19 Nov Submit your article to this journal Article views: 203 View related articles Full Terms & Conditions of access and use can be found at Download by: [ ] Date: 03 March 2017, At: 08:52

2 Engineering Applications of Computational Fluid Mechanics Vol. 4, No. 4, pp (2010) PERFORMANCE CHARACTERISTICS OF FLUSH ANGLE-OF-ATTACK MEASUREMENT SYSTEM INTEGRATED ON A PITOT TUBE J. Masud Department of Mechanical & Aerospace Engineering, Institute of Avionics & Aeronautics, Air University, E-9/4, Islamabad 44000, Pakistan Jehanzeb.Masud@mail.au.edu.pk ABSTRACT: In this paper the performance characteristics of a pressure based flush angle-of-attack measurement system, integrated on a Pitot-Static tube, is presented. The non-dimensional pressure difference between two static pressure ports located on the upper and lower surface of the conical interface, with half-cone-angle of 20, of the Pitot tube shows linear behavior (desirable) with angle-of-attack in the subsonic regime while in the supersonic regime it has non-linear behavior (undesirable). This non-linear behavior is attributed to shockwave induced flow separation just upstream of the conical interface. The undesirable non-linear trend can be eliminated by reducing the half-cone-angle of the conical interface from 20 to 12. The reduction in half-cone-angle eliminates the upstream shockwave induced flow separation in the supersonic regime thereby restoring the linear behavior between nondimensional pressure difference and angle-of-attack. However, reducing the conical interface half-cone-angle also reduces the sensitivity of the angle-of-attack measurement system in the subsonic regime by 18%. Computational Fluid Dynamics tools have been used in the analysis and refinement of flush angle-of-attack measurement system. The complete Mach number regime (Mach<2) has been computationally analyzed for the two conical interfaces. The results presented in this study have been validated by subsequent wind tunnel tests. Keywords: pitot tube, CFD, angle-of-attack 1. INTRODUCTION Pitot-Static tube (Garcy, 1980; Letko, 1947) (referred commonly as Pitot tube) is one of the important air data sensors and is designed to measure correct ambient static and total pressure corresponding to aircraft flight conditions. These pressure measurements translate into aircraft speed, pressure altitude, Mach number, vertical velocity information etc that are necessary for multiple aircraft subsystems. The measurement of correct ambient (atmospheric) static pressure by the Pitot tube over the whole Mach number regime (subsonic to supersonic) of the aircraft is the most critical aspect in its design (Garcy, 1980). Generally, a Pitot tube is placed ahead of the aircraft where the local static pressure in the subsonic regime is higher than the ambient static pressure, i.e., positive Cp value (Garcy, 1980), and is referred to as the Cp position error. In the supersonic regime the local static pressure ahead of the aircraft equals the ambient static pressure i.e. Cp position error value is zero. An aerodynamic-compensation Pitot tube is designed to compensate for the position error by generating opposite local Cp by a carefully contoured profile (Garcy, 1980; Letko, 1947; Ritchie, 1959; Lao and Bao, 1987; Latif et al., 2007; Masud and Akram, 2010; Kim et al., 2008). Typically the Pitot tube is sensitive to many factors (Leng et al., 2001) such as mounting angle (Sun et al, 2007) and pressure port shape (Sun et al., 2007) etc. For certain applications the pitot tube maybe required to house other sensors, most common amongst these are vane type Angle-of-Attack (AOA) sensor. Additionally a flush static pressure based AOA measurement system may also be placed on a Pitot tube, as was the case in our pervious study (Masud and Akram, 2010) involving design refinement of Pitot tube for supersonic aircraft. The pressure based AOA system works by measuring static pressure difference generated on the top and bottom surfaces of a cone like body at an angle-of-attack (AOA). This pressure difference increases with increasing AOA and can be used for measuring AOA in flight, thus this system is termed as the Pressure differential AOA measurement system. A representative pressure differential AOA measurement system incorporated on a pitot tube is shown in Fig. 1. The layout and design of pressure differential AOA system involves careful balance between maintaining linear (preferable) relationship between non-dimensional pressure difference and AOA that requires smaller half- Received: 15 Apr. 2010; Revised: 15 Jun. 2010; Accepted: 23 Jun

3 cone angle (HCA) and greater AOA measurement sensitivity that requires larger half-cone angles. In this paper, performance characteristics of such an AOA measurement system, integrated on a Pitot tube (Masud and Akram, 2010), is analyzed and subsequent refinement is done to overcome certain undesirable characteristics. The basic Pitot tube with the AOA conical interface (Masud and Akram, 2010) was developed as part of air data system of a supersonic fighter aircraft (Masud and Akram, 2010; Jennings, 2008). Computational Fluid Dynamics (CFD) techniques have been used in the analysis and redesign in the subsonic and supersonic regimes with basic validation of computed results done by wind tunnel test (WTT) data, which is available as part of aircraft development process. 2. COMPUTATIONAL SETUP Although the Pitot tube is symmetric about its longitudinal axis, however for AOA measurement system analysis 3D computations with a plane of symmetry are necessary. The coordinate system and other parameters of the pressure differential AOA measurement system and the Pitot tube are shown in Fig. 1. The compensation region profile is based on higher order polynomial as discussed in our earlier work (Latif et al., 2007). For CFD analysis, in addition to the Pitot tube, the complete flow domain is modeled that includes the free stream far field, the Pitot tube, and the symmetry plane. The far field is kept more than 100 times the pitot tube maximum diameter away to model free flight conditions, which is 4.5 meters from the pitot tube surface in all directions. Mapped hexahedral elements were used to mesh the computational domain. Elements were graded towards the surface of Pitot tube in order to accurately resolve the flow phenomena in its vicinity. A representative grid for the present study is shown in Fig. 2. In the present work, compressible RANS system of equations with constant property air (except density) was solved using the coupled implicit algorithm of Fluent CFD code (Fluent, 2004; Gambit, 2005). No-slip velocity boundary condition was enforced at the surfaces/ walls of Pitot tube (Fig. 1 and Fig. 2). Pressure/ velocity boundary conditions were used at stream corresponding to the desired free stream Mach number. Symmetry boundary condition was specified on the symmetry plane (Fluent, 2004). 2.1 Turbulence modeling For the present study, the Realizable k- (RKE) and the one-equation Spalart-Allmaras (SA) (Spalart and Allmaras, 1992) turbulence models were used to determine the sensitivity of computed results with the turbulence model. For the wall treatment, standard wall functions formulation of Fluent was used. Although Spalart-Allmaras is a low Reynolds number model, however for the wall treatment, standard wall functions formulation of Fluent for SA model was used. This reduces the computational cost by allowing somewhat coarser mesh near solid walls based on y +. The results in subsonic and supersonic regime indicate less then 0.5% difference in computed pressures (variable of interest for flush pressure differential angle-ofattack analysis) at the conical interface (Fig. 1) between the two (RKE and SA) turbulence models. These results were also in agreement with subsequent wind tunnel tests (discussed later in this paper). Therefore, all the computed results presented in this study are based on either the Realizable k- or the Spalart-Allmaras turbulence model with standard wall functions (Fluent, 2004). Fig. 1 Fig. 2 Geometric layout of the Pitot tube with pressure differential AOA measurement system. Computational mesh in the vicinity of the Pitot tube (oblique view, only surface mesh shown). 550

4 2.2 Grid independence analysis Numerically computed results change with the type and fineness of the mesh/ grid used for computations. In order to determine the dependence of the computational result on the mesh/ grid used, two cases with different grids (Grid 1 and Grid 2) were computed each at Mach number 0.9 with AOA 25 and Mach 1.8 with AOA 12, Grid 1 had nodes and Grid 2 had nodes. The results indicate a percentage change of less than 1.3 % and 1.8 % in the computed pressure value at the static ports locations of interest (Fig. 1) between the two grids at subsonic and supersonic speeds, respectively. The value of computed turbulent law y remains below 90 and 150 at subsonic and supersonic speeds (White, 1991; Fluent, 2004), respectively for both the grids. These values of turbulent law y show reasonable resolution of grid near the walls with respect to the usage of standard wall functions (White, 1991; Fluent, 2004). Majority of the computations for the present work were done with a mesh having grid density equivalent to Grid 2 or better. The computed results with Grid 2 are also in agreement with Wind Tunnel Test data, as shown later in this paper. 3. WIND TUNNEL TEST DATA The designed original Pitot tube shown in Fig. 1 was tested in a high speed wind tunnel as part of aircraft (Jennings, 2008) air data system development process. The high speed wind tunnel test section was 0.6x0.6 m 2 while the Pitot tube model (1:1 scale) had a maximum diameter of m; therefore wind tunnel corrections to the measured Cp were considered unnecessary and direct comparison between the computed results and the corresponding wind tunnel test data can be made. The original Pitot tube (Fig. 1) and its wind tunnel model (1:1 scale) had static pressure ports midway on the conical interface on the upper and lower surface as shown in Fig. 1. During wind tunnel testing, static pressure (and corresponding Cp) was measured and recorded for a range of Mach numbers and AOAs at these static pressure ports (sideslip was fixed at zero for these tests). All wind tunnel test (WTT). For computations, care was taken to place a node at the Pitot tube static port location so that computed results are directly available for comparison with wind tunnel test data without resorting to data interpolation. Comparison between computed and corresponding WTT data is quite satisfactory as shown later in this paper. Subsequently, wind tunnel tests were also done for similar Pitot tube but with modified AOA conical interface with shallower ( 12 ) half-cone angle (HCA). The reason for this modification and comparison of results are part of the Results and Discussion section (IV) of this paper. All these wind tunnel tests are part of aircraft (Jennings, 2008) development process as discussed earlier. 4. RESULTS AND DISCUSSION 4.1 Performance parameter of pressure differential AOA measurement system The performance characteristics of pressure differential AOA measurement system is measured in the form of derived non-dimensional pressure difference parameter K that is defined as: K=(P l -P u )/(P o -P ) where P l and P u are the pressures measured on the lower and upper surface static pressure ports and P o and P are the free stream total and static pressures, respectively. As discussed earlier, the principle of this type of AOA measurement system is that as the true AOA increases the pressure difference between the upper and its corresponding lower port (Fig. 1) increases which is represented by increasing values of K with increasing AOA. The rate of increase in K (gradient) with true AOA is a direct function of the conical interface half-coneangle (HCA). For the present Pitot tube, the designed AOA conical interface has an HCA= 20 as shown in Fig. 1. This HCA is 5 less than that reported in our earlier study (Masud and Akram, 2010). 4.2 Flow field around the basic configuration (Fig. 1, HCA=20 o ) In the present study, computations were carried out at Mach numbers of 0.22, 0.6, 0.8, 0.95, 1.1, 1.2, 1.5 and 1.8 in order to adequately cover the wind tunnel test (WTT) data range as well as the supersonic aircraft flight envelope. The AOA for the computations was set at 4, 8, 12 and 25 degrees in the subsonic regime, and 4, 8, 12 degrees in the transonic to supersonic regime. For M=1.5 additional computations were done at AOA=1 and 16 degrees as well to verify certain trend in K factor discussed later in this paper. The sideslip angle (Beta) was kept at zero for all computations. After the computations most of the post processing of the output data was done on Fluent and Tecplot software. 551

5 Fig. 3 Mach number contours on pitot tube surface and the symmetry plane for subsonic free stream (HCA=20 o ). conical interface. These figures are generally representative of flow field at all subsonic Mach numbers. The subsonic flow adjustment (slowing down) ahead of the pitot tube tip and the AOA conical interface are evident from Fig. 3. The velocity vector plot (Fig. 4) shows attached flow in the vicinity of the AOA conical interface at all AOA. However at high AOA, ( 25, M=0.8), a small region of separated flow is evident just aft of the AOA conical interface (Fig. 4). This region of separated flow is not expected to have a significant effect on the pressure or flow distribution on the AOA interface itself due to its small size and downstream location. This aspect is verified later in this paper. The supersonic flow field around the pitot tube consists of free stream flow being modified by shock and expansion waves due to the presence of the tube. Fig. 5 shows the Mach number variation on the pitot tube surface and the symmetry plane for M=1.5 and AOA= 4, while Fig. 6 shows the corresponding vector plot in the region of AOA conical interface. These figures are generally representative of flow field at all supersonic Mach numbers. The supersonic flow adjustment around the pitot tube and the AOA conical interface is characterized by three dimensional shock and expansion waves as evident from Fig. 5. The velocity vector plot (Fig. 6) shows a region of separated flow just ahead of the AOA conical interface upper surface. This is caused by interaction between shock wave and the viscous boundary layer whereby the boundary layer separates due to the strong adverse pressure gradient across the shock wave. This separated flow reattaches itself at the leading edge of the Fig. 4 Velocity vectors in the vicinity of conical interface for subsonic free stream (HCA=20 o ). The subsonic flow field around the pitot tube consists of free stream flow being modified by the presence of the tube itself. The flow adjacent to the pitot generally follows the pitot contour while far away it merges with the free stream flow. Fig. 3 shows the Mach number variation on the pitot tube surface and the symmetry plane for M=0.8 and AOA= 4, while Fig. 4 shows the corresponding vector plot in the region of AOA Fig. 5 Mach number contours on pitot tube surface and the symmetry plane for supersonic free stream (HCA=20 o ). 552

6 Fig. 6 Velocity vectors in the vicinity of conical interface for supersonic free stream (HCA=20 o ). Fig. 7 K Factor variation with AOA for subsonic/ low transonic free stream Mach numbers (HCA=20 o ). AOA conical interface and thus forms a separation bubble. One way of reducing the intensity of this shock wave and avoiding the separation bubble is by reducing the AOA interface HCA. At an AOA of 16 degrees (M=1.5), the small separation bubble is reduced in size as shown in Fig. 6 (AOA= 16 ). This region of separated flow (present at low and high AOA) is expected to have some effect on the pressure distribution on the upper surface of AOA interface due to its upstream location and may influence the corresponding K factor value. This aspect is discussed later in this paper. 4.3 Performance characteristics of AOA measurement system (HCA = 20 o ) The variation of computed K factor with true AOA is shown in Fig. 7 for subsonic and low transonic Mach numbers. The results of wind tunnel tests (WTT) are also included in this figure for comparison. The agreement between WTT results and corresponding CFD analysis is reasonable given the uncertainty caused by pitot tube manufacture tolerance and WTT data scatter (model alignment, etc.). The linear behavior of computed K factor (desirable) at subsonic Mach numbers is evident from Fig. 7. The slight variation in K factor linearity of WTT results at low transonic Mach numbers (M=0.95, Fig. 7) can be caused by a number of factors, however, within the limitations of the present 3D computational analysis, an aerodynamic reason for this small deviation could not be identified. At supersonic and high transonic Mach numbers, the variation of computed K factor with true AOA in shown in Fig. 8. The result of wind tunnel tests (WTT) is included in this figure for comparison. The agreement between WTT results and corresponding CFD analysis is reasonable given the uncertainty caused by pitot tube manufacture tolerance, model alignment and WTT data scatter. The non-linear behavior (undesirable) of computed and WTT K factor is evident from Fig. 8. The readily identifiable aerodynamic reason for this anomaly is the presence of small shock induced separation 553

7 Fig. 8 K Factor variation with AOA for supersonic/ high transonic free stream Mach numbers (HCA=20 o ). Fig. 9 Velocity vectors in the vicinity of conical interface for supersonic free stream (HCA=12 o ). bubble on the upper surface at the forward junction of the AOA conical interface as discussed earlier (Fig. 6). This bubble modifies the external flow in its vicinity thereby modifying the pressure at the upper port of the AOA measurement system, which in turn causes the K factor value to vary from linear behavior. The upstream location of the separation bubble aggravates the non-linear phenomenon (Fig. 6). The AOA conical interface half-cone angle (HCA) plays a major part in the formation of the separation bubble at supersonic speeds since the strength of the conical shock wave in front of it (and subsequent shock wave interaction with boundary layer) is dictated by it (Schreier, 1982). Reducing the AOA conical interface half-cone angle can reduce or even eliminate the separation bubble and all its detrimental effects on K factor, as is discussed next. 4.4 Design refinement with reduced halfcone-angle (12 o ) AOA interface The non-linear behavior of K factor for AOA interface with HCA= 20 in the supersonic regime presents multiple challenges before the AOA information from it can be used for the aircraft air data system. Due to this fact the air data system designers decided to reduce the HCA of AOA interface from 20 to 12 for the Pitot tube. For this modified AOA conical interface, additional computational analysis was done. For this analysis, the existing 3D geometry and computational domain was modified and remeshed in the AOA interface region. The rest of the geometry, constraints and conditions were not changed. Emphasis of this analysis is placed on supersonic regime where the nonlinear K behavior is apparent (Fig. 8), while subsonic regime is also analyzed for verification. The results of this additional analysis are discussed next. Fig. 9 shows the velocity vector plot for the modified AOA interface for M=1.5 at 4 deg AOA. It is evident that the shock wave induced separation bubble near the AOA interface is not present in this case. The side-by-side comparison of 20 deg HCA and 12 deg HCA AOA interface for the Pitot tube is shown in Fig. 10 for same supersonic (M=1.5, AOA= 4 ) and subsonic (M=0.8, AOA= 4 ) conditions. The absence of the small separation bubble at supersonic speeds for the 12 deg HCA interface is also apparent from Fig. 10. The completely attached flow in the vicinity of the AOA interface (12 deg HCA) is due to weaker conical shock wave (Schreier, 1982; White, 1991) created by the smaller angle of the interface, which in turn does not create sufficiently large adverse pressure gradient to locally separate the boundary layer. Ahead of the AOA interface, the Mach number distribution (and corresponding flow and pressure distribution) is the same for 20 deg HCA as well as 12 deg HCA AOA interface for M=1.5. This is due to the supersonic nature of the flow where the presence 554

8 of different AOA interfaces cannot be sensed upstream. The side-by-side comparison of 20 deg HCA and 12 deg HCA AOA interface under subsonic conditions (M=0.8, AOA=4 deg, Fig. 10) shows attached flow in the vicinity of both AOA interfaces, which is signified by the regular pattern of the Mach number distribution. This is also verified by the velocity vector plot (Fig. 4) discussed earlier. Immediately ahead of the AOA interface, the subsonic Mach number distribution (and corresponding flow and pressure distribution) is slightly different for 20 deg HCA and 12 deg HCA AOA interface. This is due to the subsonic nature of the flow where the presence of different AOA interfaces can be sensed upstream of the flow. However in the compensation region the effect of change in downstream AOA interface (from 20 deg to 12 deg HCA) is not distinguishable from Fig. 10 (due to the scale) but is nevertheless present to certain degree as identified in our earlier work (Masud and Akram, 2010). The variation of computed K factor with true AOA for Pitot tube with reduced HCA (12 deg) AOA interface is shown in Fig. 11 for subsonic and supersonic Mach numbers. The results of a subsequent wind tunnel test (WTT) on similar configuration, as discussed earlier, are also included for comparison. The agreement between WTT results and corresponding CFD analysis is reasonable given the uncertainty caused by Pitot tube wind tunnel model manufacturing tolerance and WTT data scatter due to various factors (model alignment etc). The linear behavior of computed K factor (desirable) in the subsonic and supersonic flow regimes is apparent from Fig. 11. This is due to the absence of undesirable aerodynamic phenomenon like localized upstream flow separation in the vicinity of the modified AOA interface (reduced 12 HCA). Therefore, it is seen that reducing the AOA interface HCA from 20 deg to 12 deg solves the K factor nonlinearity problem at supersonic speeds as encountered earlier. However, there is a downside to this reduction in AOA interface half-cone angle (HCA) as is discussed next. Reducing the AOA interface HCA generally reduces the value of K/ AOA i.e. the gradient Fig. 10 Comparison of effect of conical interface half-angle (HCA) on upper surface Mach number distribution. Fig. 11 K Factor variation with AOA for subsonic/ supersonic free stream Mach numbers (HCA=12 o ). 555

9 of the curve on K versus AOA plot (refer Figs. 7, 8 and 11). Due to non-linearity of K in supersonic regime for HCA=20 deg AOA interface (Fig. 8), the effect of reduction in HCA (from 20 to 12 deg) on K/ AOA cannot be assigned a single quantitative value for comparison. However, in the subsonic regime both 20 deg and 12 deg HCA AOA interfaces produce generally linear behavior (Figs. 7 and 11) and the reduction in HCA from 20 deg to 12 deg results in reduction in K/ AOA from to per deg, which is about 18 % reduction. This represents a reduction in effectiveness or sensitivity of the AOA measurement systems for similar pressure measuring capability (pressure sensors etc), which is a design compromise for the present Pitot tube. 5. CONCLUSION It has been computationally shown (and experimentally verified) that the pressure differential angle-of-attack measurement system mounted on a Pitot tube conical interface with half-cone angle of 20 produces an undesirable non-linear behavior between non-dimensional pressure difference K and angle-of-attack (AOA) in supersonic regime, which is due to the presence of an upstream shock-induced separation bubble. Reducing the half-cone-angle of the angle-of-attack interface from 20 degrees to 12 degrees resolves this problem and the desirable linear behavior in both subsonic and supersonic flow regimes is achieved, but at a reduced AOA measurement effectiveness/ sensitivity. NOMENCLATURE AOA Angle-of-Attack (Beta) Side slip angle CFD Computational Fluid Dynamics Cp 2 Pressure coefficient (P-P )/½ V HCA Half Cone Angle K Non-dimensional pressure difference K=(P l -P u )/(P o -P ) M Free stream Mach number P Static pressure P o Total pressure P l Lower surface static pressure P u Upper surface static pressure P Free stream static pressure RKE Realizable K-є turbulence model SA Spalart-Allmaras turbulence model V Free stream velocity Free stream density WTT x y y + REFERENCES Wind Tunnel Test Axial coordinate Radial coordinate Non-dimensional length scale associated with turbulence model 1. FLUENT (2004). Computational fluid dynamics software package, Ver , Fluent Inc, Lebanon, NH, USA. 2. FLUENT (2004). Computational fluid dynamics software package user guide, Ver , Fluent Inc, Lebanon, NH, USA. 3. GAMBIT (2005). Geometry and mesh generation software package, Ver Fluent Inc, Lebanon, NH, USA. 4. Garcy W (1980). Measurement of aircraft speed and altitude. NASA Reference Publication Jennings G (2008). JF-17 Production commences. Jane s Defense Weekly 45(5): Kim DJ, Cheon, YS, Myong RS, Park CW, Cho TH, Park YM, Choi IH (2008). Design of Pitot-tube configuration using CFD analysis and optimization techniques. Transactions of the Korean Society of Mechanical Engineers, B 32(5): Lao S, Bao Y (1987). Computation of transonic aerodynamically compensating pitot tube. AIAA Journal of Aircraft 25(6): Latif A, Masud J, Sheikh SR, Parvez K (2007). Robust design of a aerodynamic compensation pitot-static tube for supersonic aircraft. Journal of Aircraft 44(1): Leng X, Zhang X, Xie J, He F (2001). Effects of Pitot tube on the measurement of supersonic flow. Qinghua Daxue Xuebao/ Journal of Tsinghua University 41(11): Letko W (1947). Investigation of the fuselage interference on a pitot-static tube extending forward from the nose of the fuselage. NACA Technical Note NACA-TN Masud J, Akram F (2010). Adjustment of aerodynamic compensation characteristics of a Pitot tube by rear-body shape manipulation. ASME Journal of Fluids Engineering 132(034502): Ritchie VS (1959). Several methods for aerodynamic reduction of static-pressure sensing errors for aircraft at subsonic, nearsonic, and low supersonic speeds. NASA Technical Report NASA TR R

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