Position Error Studies for Air Data Boom of Flying Wing Configuration using RANS CFD Simulation

Transcription

1 20 th Annual CFD Symposium August 9-10, 2018, NAL, Bangalore, India Position Error Studies for Air Data Boom of Flying Wing Configuration using RANS CFD Simulation G Prabhakaran*, Arvind Bobby Alphonse, B Saravanan, K Selvaraj ῼ Aerodynamics Division Aeronautical Development Establishment, Bangalore ABSTRACT: Accuracy in air speed measurements is critical for safe operation and optimal performance of the aircraft. If the measured air speed is less than the actual speed, aircraft may enter the stall regime under certain conditions. On the other hand if the measured air speed is more, then performance of the aircraft will be affected. Air speeds are measured by air data boom located ahead of the aircraft. The static pressure and total pressure measured by air data boom are translated into air speed and altitude. The static pressure data measured by air data boom are subject to instrument errors and position errors. Instruments error are due to errors in the standalone probe, plumbing errors and errors associated with calibration which are typically very small. The major portion of the error is the position error which is due to the location of static pressure port. This position error depends on the configuration of aircraft and the stand-off distance from nose of the aircraft. In this paper the effect of flying wing airframe on the position error is presented. A commercially available air data boom with pitot static probe, angle of attack vane and angle of side slip vane is opted for analysis. Studies are carried out for two different boom lengths and the effect of increase in standoff distance is brought out. This work also presents the computational fluid dynamics requirements and methodologies to estimate the position error for alpha vane, beta vane and static probe, which are located on the boom. The analysis is carried out at M=0.2 for the angle of attack range from -8 to 16.The measured velocity is less than the free stream velocity as expected in the lower range of angle of attack and it decreases with increasing angle of attack. At higher angle of attack, the probe measures a higher velocity as compared to the free stream due to local flow acceleration around the boom. Based on the computational results, the error in velocity and flow angles are quantified. Keywords: Alpha and Beta Vane, Air data Boom, Position Error, Flying Wing Nomenclature IRA = Indian Reference Atmosphere. AOA = Angle of Attack in degree. AOSS = Angle of Side Slip in degree. V = Free stream velocity in m/s. P o = Free stream total pressure in Pascal. P = Pressure Measured by static probe in Pascal. a = Speed of sound in m/s. V measured = Velocity measured by Vane in m/s; V measured = a ( 2 γ 1 ) γ 1 [(P o ) γ 1] P ρ = Free stream density in Kg/m 3. V error = Velocity error in Percentage; [(V measured -V ) X100 ] / (V ) P = Free Stream Static Pressure in Pascal. V x = Velocity in X-direction in m/s. V Y = Velocity in Y-direction in m/s. *Research Fellow Scientist Scientist ῼ Scientist

2 V z = Velocity in Z-direction in m/s. α vane = AOA indicated by AOA vane; tan -1 (V Z/V X). α actual = Free stream AOA. α error = AOA error; α vane - α actual vane = AOSS indicated by AOSS vane; tan -1 (V Y/V X) actual = Free Stream AOSS. error = AOSS error; vane - actual. C P = Pressure Coefficient; (P-P )/(0.5*ρ * V * V ) 1. Introduction: Pitot-static tube is an important air data sensor and is designed to measure the ambient static pressure and total pressure corresponding to the aircraft flight conditions. The pressure measured by tube is translated into aircraft speed, altitude, Mach number, vertical velocity information etc. that is necessary for various aircraft subsystems. The measurement of correct ambient pressure by the Pitot tube over the whole Mach number of aircraft is critical 1. Commonly, Pitot tube is placed ahead of the aircraft where the local static pressure in the subsonic flight condition is higher than the ambient static pressure, i.e. positive Cp value 1. The presence of the aircraft in the air stream causes error to the measuring instruments. The local flow angles also differ from the free-stream flow direction. Accurate measurements of flow angles, static pressure, and total pressure are necessary for safe flight operations. Several flight subsystems, such as auto flight controls, engine controls, and cockpit and cabin environmental control depend on accurate air data. References 5 and 6 specify the accuracy level required of pitot static tube by civil and military organizations. Flow angle vanes are small, mass balanced winglets designed to align with free stream of air during flight and it is the most common method for measuring the angle of attack and side slip in civil aviation. If the vane is fitted ahead of the airframe, it will allow the vanes to read the angles near to the free stream conditions. The vane is mounted behind the airframe and is prone to more error because of the influence of wing and propeller etc. Ideally, a system measuring flow angle such as vanes and differential pressure sensors or probes, should be fitted to a boom usually extended out from the aircraft i.e well away from influence of airframe on the flow field 2. All systems for measuring angle of attack and sideslip mentioned above are susceptible to error due to a number of factors 3. In this paper, performance characteristics of flow vanes and static probe are integrated on the air data booms and it is analyzed for various flight conditions. There are two types of boom configurations used for analysis, one is called as an original boom and another one is an extended boom (265mm longer than original boom). Computational Fluid dynamics (CFD) techniques have been used in the analysis to estimate the recommended extension length of the nose boom which is to be installed in the first flight of flying wing. 2. Computational Methodology: Commercial CFD software ICEM CFD, HiFUN and Tecplot are used as Pre-processor, Solver and Post processor respectively. The details of Pre-Processor, solver and post-processor are given below 2.1 Preprocessing The geometry is cleaned using ICEM CFD software. The air data boom is modeled without vanes as they will have negligible influence over the local flow field. Spherical computational domain is used for generating volume mesh around the model. The computational domain radius is taken as 10 times of the fuselage length. The mesh size of 17 million is used to capture all the geometry features and to keep the wall Y + value below 5. The typical grid has been shown in the Fig.1.1 and 1.2.

3 Fig.1.1 Flying Wing with Computational Domain Fig.1.2 Nose Boom 2.2 Solution The flow solver is a Reynolds Averaged NS (RANS) equation solver. Second order discretization scheme is used to solve for all flow variables. Inviscid and viscous flux calculation is carried out using Roe and Green Gauss schemes respectively. The Spalart Allmaras turbulence model is used for turbulent viscosity calculations. IRA sea level conditions are assumed for CFD analysis. Pressure far-field conditions have been used for the computational domain boundary. No-slip adiabatic wall condition is applied to all solid surfaces. The solution was deemed converged when the aerodynamic coefficients were changing less than 0.5% with iteration. The Free stream conditions and reference values used in the solver are given in Table 1. Table 1: Free stream Conditions and references Values Parameter Mach Number 0.2 Values Density(kg/m 3 ) Pressure(N/m 2 ) Temperature(K) Viscosity (Kg/m-s) X 10-5 Ratio of Specific Heats Post-Processing Tecplot is used for post processing the results of HiFUN solver. Data have been analyzed and presented in the subsequent section. The velocity components are measured in three directions using probe option and average pressure is calculated around static pressure probe. In Fig.2, the locations of alpha vane, beta vane, and static probe are shown.

4 Location of Alpha Vane Location of Static Pressure Probe Location of Beta Vane Fig.2 location of Vanes and Static pressure Probe 3. Boom Geometry Description The air data boom with pitot static probe is shown in Fig-3. Fig-3: Straight Nose Air Data Boom-Heated 4. Results and Discussion Because of the flow field created by the flying wing, the flow angle at any given location in the vicinity of the flying wing will generally differ from the true angle of attack of flying wing. At subsonic speeds the effects of the flow field extend in all directions. As the flow field around the each configuration is unique, in order to analyze the CFD results, first a vane center point was defined in Tecplot, then three velocity components are calculated at this point. The Computational simulation matrix is given in Table 3. Table 3: Summary of CFD Simulation matrix Parameter Values Mach Number 0.2 Angle of attack (deg) -8 to 16 in steps of 2 Beta(deg) 2 to 10 in steps of 2

5 The static pressure is measured in two ways during flight, with Pitot-static tube or static probe on the fuselage. The static pressure distribution on the air data boom is influenced by various factors like aircraft angle of attack, sideslip angle, speed etc. The surface distribution of the static pressure error (ΔP/q ) along the boom length is used to reveal the local variations due to major elements of aircraft, control surfaces, and antenna (see Fig.4).There is a gradual decrease in velocity error upto 10degree. Due to favorable pressure gradients around the air data boom there is an increase in error at higher angle of attack, which is shown in Fig-9. Favorable pressure gradients are observed along the boom length, at 12degree angle of attack due to local flow acceleration as shown in Fig-5 and 6. From Fig-10, it is observed that velocity error reduces with increasing sideslip. The variation of pressure coefficient across the boom at various locations in longitudinal direction for both booms are shown in Fig-7 and 8. It is observed that favorable pressure gradient increases along the boom length from the aircraft nose. There will be difference in local AOA measured by the vane and free stream AOA due to the location of the AOA vane. The local streamline direction is estimated by estimating the local velocity components. It follows from Fig-11 and Fig-12 that angle of attack error increases with angle of attack and angle of sideslip error increases with sideslip. In all cases, it is observed that extended boom is associated with lesser position error Aircraft Nose ΔP/q X/D Fig-4: CFD simulated static pressure-error (ΔP/q ) distribution on the original boom plotted against boom centerline axis

6 Fig-5: Static Pressure Variation along the original boom (Side View) Fig-6: Static Pressure Variation along the extended boom (Side View) Fig-7: C P Variation across the original boom from aircraft nose for12deg angle of attack (Front View)

7 V error (%) Fig-8: C P Variation across the extended boom from aircraft nose for 12deg angle of attack (Front View) Extended Boom Original Boom AOA (deg) Fig-9: Variation of V error with AOA

8 AOA error (deg) V error (%) Extended Boom Original Boom AOSS (deg) Fig-10: Variation of V error with AOSS Extended Boom Original Boom AOA (deg) Fig-11: Variation of AOA error with free stream AOA

9 AOSS error (deg) Extended Boom Original Boom AOSS (deg) Fig-12: Variation of AOSS error with free stream AOSS 5. Conclusion: There is a gradual decrease in velocity error upto 10deg for both the booms, the measured velocity is said to lesser than free stream velocity. An increase in velocity error is observed after 10 degree due to local acceleration around the boom, the measured velocity is said to higher than free stream velocity. Different length of air data boom are taken to estimate the error values such as velocity error, angle of attack and angle of sideslip error. Marginal reducation in error can be seen when the length of the boom is increased. The effect of flying wing airframe on air data boom is less significant when compared to other configurations.

10 6. References 1. Gracy,W., Measurement of Aircraft Speed and Altitude, NASA Reference Publication 1046, G.M Sakamoto. Aerodynamic characteristics of vane flow angularity sensor system capable of measuring flight path accelerations for Mach number range from 0.40 to NASA technical Note,(TN D_8242), C.J.Bennett, N.J.Lawson, J.E. Gautery, A.Cooke. CFD simulation of flow around angle of attack and sideslip angle vanes on a bae jetstream 3102-part.Submitted to aerospace Science and technology, Gracey, W., Measurement of Static Pressure on Aircraft, NACA Rept. 5. Wusest, W., Pressure and Flow Measurement, AGARD AG-160, Vo1.11, AGARD Flight Test Instrumentation Series, Federal Aviation Administration, Federal Aviation Regulations, Volume 3,Transmittal m11,part25, Airworthiness standrads: Transport Category Airplanes U.S. Government Printing office,washington,dc,june U.S.Nay, Bureau of Aeronautics, Instrument Systems, Pitot Tube and Flush Static Port Operated, Installation of, MIL-I-6115A,Mar Haering, E.A., Airdata Measurement and calibration, NASA TM , Installation and Inspection of Pitot-Static Pressure Systems U.S. Air Force MIL-P_26292 MIL-P-26292C, Wright-Patterson AFB, J.A Lawford and K.R Nippress., Calibration of air-data systems and flow direction sensors. Advisory Group for Aerospace Research and Development flight test techniques series (AG-300), J.D.Anderson. Computation Fluid Dynamics-First Edition McGraw-Hill Education, New York, USA, 1995.