Statistical analysis of CFD results for missile surface pressures

Accurate prediction of complex rotor-airframe integration is essential for optimum design and accurate analysis of rotorcraft. Over the years, a broad array of computational methodologies has been employed to pre dict the rotor-airframe interaction. In the present study three leading computational approaches are developed and employed in the study of rotor -airframe integration problems. These approaches include an overset grid -based, unstructured tetrahedral -based, and unstructured Cartesian grid - based methods. The computational results are compared with experimental data on a generic rotor - airframe configuration tested at Georgia Tech and on the Robin helicopter geometry tested at NASA Langley. The agreemen t with experimentally measured surface pressure is found to depend strongly on the assumed rotor load distribution . These studies are valuable in helping develop and document the capability of these approaches to provide accurate predictions of rotor -airframe interaction. Introduction The flowfield around a single rotor helicopter, whether in forward flight or hover, is difficult to model due to the presence of strong vorticity which lingers in the vicinity of the rotor. In modern helicopter designs, the rotor blades are placed very close to the fuselage for compact configurations and to reduce the drag associated with the mast. In such configurations, the vortex system generated by the rotor blades interacts with the body surface in a complex unsteady manner. Rotors and airframes designed without taking into account the rotor -airframe aerodynamic interactions tend to perform significantly below their predicted capabilities. 1 Over the years, a broad array of computational methods have been employed to predict the rotor -airframe interaction. Several reviews of these methods are available. 2,3,4 These methods include modeling of individual rotors and tip vortices and modeling the rotor as a single actuator disk.

Computational and experimental results are presented for delta wings with vortex flaps. The wings have an undeflected leading-edge sweep of 75 deg. Flap angles of 5 deg and 10 deg, measured in the streamwise direction, are considered. The nominal angle of attack alpha is varied from 4 deg to 12 deg. Results for freestream Mach numbers of 1.7 and 2.4 are shown. Surface pressure, tuft patterns and vapor screens are given for the experimental data. Surface pressure, tuft patterns, cross-flow velocities, total pressure loss and cross-flow Mach number are given for the numerical data. C(l)-vs-alpha curves are shown for experimental and computational results. The flows are shown to be very sensitive to angle of attack, and the agreement between experimental and theoretical results is improved if the calculations are made at angles of attack slightly larger than the experimental angles of attack. The computational model correctly predicts the topology of the flow in each of the cases considered. The lift is predicted well at the higher angles of attack, but slightly overpredicted at the lower angles of attack.

Research into autonomous or remotely piloted vehicles has stimulated interest in the CFD analysis of these types of aircraft designed around non-traditional mission requirements. As a joint CFD code validation effort by The Technical Cooperation Program (TTCP), a consortium of governmental interests in five countries, a study is being conducted into the performance predictions of various Boeing/US AFRL model 1303 UCAV configurations with different leading edges. The purpose of this study is to assess the current CFD capabilities for UCAV type vehicles of the TTCP countries in order to identify shortcomings and through this collaborative effort improve the performance of the codes. This multi-year study includes a baseline, as well as sharp and rounded shaped profiles. Angles-of-attack range from -5 to 25 degrees. A variety of government and commercial flow solvers have been applied, including AVUS, COBALT, OVERFLOW, NPARC, and FLUENT. Various turbulence models are also examined and included in this effort. CFD results have been compared using lift, drag, moment, and pressure coefficient profiles to experimental data, collected at the QinetiQ low-speed 5 meter wind tunnel in Farnborough, U.K., as well as code to code comparisons. After the first year of results was obtained using a baseline leading edge, it was seen that both similarities and differences exist among the results from the various CFD codes and with experiment. Analysis of the data reveals that various strengths and weaknesses in the CFD approaches influence the agreement with experiment. This comparative study offers the opportunity to leverage and improve CFD approaches among the various participants in order to attain improved computational capabilities for planned future efforts.

Grid adaptation using general elements was applied to a hypersonic high angle of attack re-entry vehicle to determine if there could be improvements made in the prediction of heat transfer starting from a baseline unstructured grid. Grid adaptation was able to greatly reduce the errors in heat transfer prediction that are common when computing on unstructured grids. The reduction in error is due to a shrinking of average cell size near the bow shock. The smaller cells cause a less drastic jump in the solution of the Riemann problem at the shock. Additionally, a hybrid numerical scheme was used to add numerical dissipation near the bow shock region to help further reduce the jump in the solution of the Riemann problem at the shock. Average errors in heat transfer near the stagnation region were reduced by approximately 50% due to the smaller average cell size and hybrid numerical scheme. Overall the approach of combining grid adaptation and a hybrid numerical scheme appears to provide a robust and reliable method of computing heat transfer rates for hypersonic re-entry vehicles using unstructured grids.

A combined experimental and computational study of pressure-driven three-dimensional separation in a turbulent boundary layer

An implicit finite-difference solver for either the Euler equations or the thin-layer Navier-Stokes equations was used to calculate a transonic flow over the NACA 64A010 airfoil pitching about its one-quarter chord. An unsteady automatic grid-generation procedure that will improve significantly the computational efficiency of various unsteady flow problems is described. The calculated results for both inviscid and viscous flows at Much number 0.8 over the airfoil oscillating with reduced frequency referenced to one-half chord, 0.2, are compared with experimental data measured in the Ames 11 x 11-ft Transonic Wind Tunnel. Nonlinear, unsteady effects of the flow on the surface pressure variations, shock-wave excursions, and overall airloads are examined. Good agreements between the results of computations and experiments were obtained. In the shock-wave region, however, the results of the viscous-flow computations showed closer agreement with the experimental data.

Investigation and simulation of a cross-flow air classifier

This paper presents an efficient numerical method for the flows past airfoils at very low Reynolds numbers, which are of interest for the micro-aircraft applications and the unmanned air vehicles (UAV). The flows at very low Reynolds numbers are dominated by viscous effects and by very thick boundary layers, with few computational or experimental results available for airfoils. The present analysis is based on a pseudo-time integration method using artificial compressibility to accurately solve the Navier-Stokes equations. This is done in a rectangular computational domain obtained by a coordinate transformation fiom the physical flow domain around the airfoil. The method uses a central differencing scheme on a stretched staggered grid in the computational domain. This method is first successfully validated for the flows with multiple separation regions past a downstream-facing step by comparison with previous experimental and computational results at very low Reynolds numbers between 400 and 1200. Then the method is used to obtain solutions for several NACA airfoils at very low Reynolds numbers between 400 and 6000. The present airfoil solutions are validated by comparison with the numerical results obtained by Kunz & Kroo for Reynolds numbers between 1000 and 6000 (no results were available for Reynolds numbers smaller than 1000). A very good agreement has been found between the two sets of results. The decrease in the low Reynolds number was found to lead to a marked increase in the pressure coefficient on the airfoil, which is more pronounced towards the trailing edge and for thinner airfoils.

Emerging and disruptive technologies are those technologies which are deemed to have potentially disruptive effects to our current way of thinking and on current Defence Systems of Systems (SoS) solutions. The Emerging and Disruptive Technologies Action Group (EDTAG) within The Technical Cooperation Program (TTCP) has successfully influenced international and national strategic planning capabilities by developing and applying an end-to-end process to identify, assess, contextualise, and report on the impact of emerging and disruptive technologies on national strategic and force development programs. The EDTAG process is widely applicable across international organisations as well as national science and technology enterprises. EDTAG validated and implemented the process, which resulted in significant insight into several new emerging science and technology areas. In this paper, we discuss the end-to-end process that EDTAG developed from 2011-2013, the activities that comprise that process, and the recommendations and lessons learned from development and implementation.

Extreme stress and adversity due to combat operations and disaster assistance is commonplace in the military. Adverse events may be compounded by multiple deployments and family or relational problems. The military’s mission readiness depends largely on the resilience of service members, family members, military units, and communities. In general, resilience is an individual, family, organization, or communities’ ability to adapt to adverse stress effectively as measured by performance and well-being. A resilience defi nition provided by The Technical Cooperation Program (TTCP) The Technical Panel 13 (TTP-13) is “the sum total of psychological processes that permit individuals to maintain or return to previous levels of well-being and functioning in response to adversity.” 1 This article provides an overview of some of the organizations that focus on resilience building for military members, providers, units, families, and communities.

ow eld in an unstructured legacy RANS computational uid dynamics code. The hybrid method consists of a blending of the k ! SST RANS model with a one-equation LES model for the subgrid-scale turbulent kinetic energy (k sgs ). Unstructured grids provide better resolution of complex geometries which is the motivation for extending this method. Correlations include theoretical data, experimental data and computational results with RANS turbulence models.

Numerical predictions of the longitudinal aerodynamic characteristics for the Ares I class of vehicles, along with the associated error estimate derived from an iterative convergence grid refinement, are presented. Computational results are based on the unstructured grid, Reynolds-averaged Navier-Stokes flow solver USM3D, with an assumption that the flow is fully turbulent over the entire vehicle. This effort was designed to complement the prior computational activities conducted over the past five years in support of the Ares I Project with the emphasis on the vehicle s last design cycle designated as the A106 configuration. Due to a lack of flight data for this particular design s outer mold line, the initial vehicle s aerodynamic predictions and the associated error estimates were first assessed and validated against the available experimental data at representative wind tunnel flow conditions pertinent to the ascent phase of the trajectory without including any propulsion effects. Subsequently, the established procedures were then applied to obtain the longitudinal aerodynamic predictions at the selected flight flow conditions. Sample computed results and the correlations with the experimental measurements are presented. In addition, the present analysis includes the relevant data to highlight the balance between the prediction accuracy against the grid size and, thus, the corresponding computer resource requirements for the computations at both wind tunnel and flight flow conditions. NOTE: Some details have been removed from selected plots and figures in compliance with the sensitive but unclassified (SBU) restrictions. However, the content still conveys the merits of the technical approach and the relevant results.

Application of a two-equation compressible turbulence model for practical hypersonic flows 18 presented. The solution algorithm is based on solving all of the governing equations simultaneously. Application of the solution procedure to several test cases for compressible flows show good agreement with theoretical predictions and/or other computational results. The solution procedure is employed to investigate the effects of turbulence in Type III and Type IV shock-shock interactions in hypersonic flows in association with adaptive unstructured grids. Computational results for these cases are compared with available experimental data.

Based on unstructured hybrid grid and dynamic overlapping grid technique, numerical simulations of Unsteady Reynolds Averaged Navier-Stokes equations were performed and investigation on isolated propeller aerodynamic characteristics and effects of propeller slipstream on turboprops were undertaken. The computational grid consisted of rotational subzone of propeller and stationary major-zone of aircraft, and walls criterion was used in the automatic hole-cutting procedure. Distance weight interpolation and tri-linear interpolation were developed to transfer information between the rotational and stationary subzones. The boundaries of overlapping grids were optimized for fixed axis rotation. The governing equations were solved by dual-time method and Lower Upper-Symmetric Gauss-Seidel method. The method and grid technique were verified by isolated propeller configuration and the computational results were in well agreement with the experimental data. The grid independence was studied to establish the numerical results. Finally, the flow around a turboprop case was simulated and the influence of propeller slipstream was presented by analyzing the surface pressure contours, profile pressure distribution, vorticity contours and profile streamline. It's indicated that the slipstream accelerates and rotates the free stream flow, changing the local angle of attack, enhancing the downwash effects, affecting the pressure distribution on wing and horizontal tail, as well as increasing the drag coefficient, pitching moment coefficient and the slope of lift coefficient.

Numerical prediction of aerodynamic characteristics of prismatic cylinder by finite element method with Spalart–Allmaras turbulence model