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The high-fidelity and reduced order CFD method's validation and analysis are presented for the Multirotor Test Bed. The reduced-order method uses a rotor-disk model in the OVERFLOW CFD code and is an important aspect of this study. A rotor-trim option loosely coupling the Comprehensive code CAMRAD II with the OVERFLOW CFD code is included in the simulation that captures rotor blade aerodynamics, blade and rotor performance, rotor-rotor, and rotor-body interactions. Detailed comparison and validation were performed for both higher-fidelity full rotor unsteady flows and reduced-order rotor-disk methods. The tunnel geometry is included in the detailed computational effort. Computations of the low-Reynolds number flows were performed for a different number of in-plane rotors as well as for a single rotor modeled inside the tunnel.
Equations and procedures are presented for numerically determining the aerodynamic characteristics of lifting rotors. The equations are general and can account for stall and compressibility effects, as well as variations in hub and blade configurations, that are normally omitted from conventional analytical rotor treatments. The application of the method to solution by automatic computing machines is discussed.
A theoretical investigation of dynamic response characteristics of rotor blades was carried out with special emphasis on torsional degrees-of-freedom. Coupled equations of motion for flapwise bending and torsion were formulated at varying azimuth positions for rotor blades with noncollinear aerodynamic, elastic and mass axes. Both structural and aerodynamic mass, damping and stiffness coefficients were included. The variations of a sample blade at different flight conditions were investigated from these equations. The obtained numerical results were illustrated. The sensitivity of overall blade vibration characteristics to torsional oscillations was also investigated from the equations of motion for the sample blade. The illustrated results show the importance of torsional degrees-of-freedom in rotor blade analysis. Various possibilities of improving the overall response by tuning blade geometric, structural and aerodynamic characteristics are discussed.
An experimental investigation of rotor/wing aerodynamic interactions in hover is described. The investigation consisted of both a large-scale and a small-scale test. A 0.658-scale V-22 rotor and wing was used in the large-scale test. Wing download, wing surface pressure, rotor performance, and rotor downwash data from the large-scale test are presented. A small-scale experiment was conducted to determine how changes in the rotor/wing geometry affected the aerodynamic interactions. These geometry variations included the distance between the rotor and wing, wing incidence angle, wing flap angle, rotor rotation direction, and configurations both with the rotor axis at the tip of the wing (tilt rotor configuration) and with the rotor axis at the center of the wing (compound helicopter configuration).
Advanced rotorcraft technology and tilt rotor aircraft were discussed. Rotorcraft performance, acoustics, and vibrations were discussed, as was the use of composite materials in rotorcraft structures. Rotorcraft aerodynamics, specifically the aerodynamic phenomena of a rotating and the aerodynamics of fuselages, was discussed.
Experimental aerodynamic performance characteristics of rotor entry vehicle configuration - subsonic
Experimental aerodynamic performance characteristics of rotor entry vehicle configuration - transonic
Experimental aerodynamic performance characteristics of rotor entry vehicle configuration - supersonic
An aeroelastic model is developed to examine the use of splitter blades as a passive detuning mechanism for flow induced forced response of unstalled supersonic turbomachine rotors. The splitters introduce aerodynamic and structural detuning to the rotor design. The relationship between aerodynamic and structural detuning and the location and chord lengths of splitters is analyzed. The model is applied to the flow induced response of four 12-blade rotors with Verdon's (1973) Cascade B flow geometry. The data reveal that for gusts characterized by forward and backward traveling waves the splitters generally decrease the maximum amplitudes of response; however, for some gust load interblade phase angles, such as -180 deg and 120 deg the splitters did not reduce the amplitudes of response.
An enhanced integrated aerodynamic load/dynamic optimization procedure is developed to minimize vibratory root shears and moments. The optimization is formulated with 4/rev vertical and 3/rev inplane shears at the blade root as objective functions and constraints, and 4/rev lagging moment. Constraints are also imposed on blade natural frequencies, weight, autorotational inertia, contrifugal stress, and rotor thrust. The Global Criteria Approach is used for formulating the multi-objective optimization. Design variables include spanwise distributions of bending stiffnesses, torsional stiffness, nonstructural mass, chord, radius of gyration, and blade taper ratio. The program CAMRAD is coupled with an optimizer, which consists of the program CONMIN and an approximate analysis, to obtain optimum designs. The optimization procedure is applied to an advanced rotor as a reference design. Optimum blade designs, obtained with and without a constraint on the rotor thrust, are presented and are compared to the reference blade. Substantial reductions are obtained in the vibratory root forces and moments. As a byproduct, improvements are also found in some performance parameters, such as total power required, which were not considered during optimization.
Various new technologies currently under development may enable controlled blade shape variability, or so-called blade morphing, to be practically employed in aircraft engine fans and compressors in the foreseeable future. The current study is a relatively brief, preliminary computational fluid dynamics investigation aimed at partially demonstrating and quantifying the aerodynamic potential of fan rotor blade morphing. The investigation is intended to provide information useful for near-term planning, as well as aerodynamic solution data sets that can be subsequently analyzed using advanced acoustic diagnostic tools, for the purpose of making fan noise comparisons. Two existing fan system models serve as baselines for the investigation: the Advanced Ducted Propulsor fan with a design tip speed of 806 ft/sec and a pressure ratio of 1.294, and the Source Diagnostic Test fan with a design tip speed of 1215 ft/sec and a pressure ratio of 1.470. Both are 22-in. sub-scale, low-noise research fan/nacelle models that have undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. The study, restricted to fan rotor blade morphing only, involves a fairly simple blade morphing technique. Specifically, spanwise-linear variations in rotor blade-section setting angle are applied to alter the blade shape; that is, the blade is linearly retwisted from hub to tip. Aerodynamic performance comparisons are made between morphed-blade and corresponding baseline configurations on the basis of equal fan system thrust, where rotor rotational speed for the morphed-blade fan is varied to change the thrust level for that configuration. The results of the investigation confirm that rotor blade morphing could be a useful technology, with the potential to enable significant improvements in fan aerodynamic performance. Even though the study is very limited in scope and confined to simple geometric perturbations of two existing fan systems, the aerodynamic effectiveness of blade morphing is demonstrated by the configurations analyzed. In particular, for the Advanced Ducted Propulsor fan it is demonstrated that the performance levels of the original variable-pitch baseline design can be achieved using blade morphing instead of variable pitch, and for the Source Diagnostic Test fan the performance at important off-design operating points is substantially increased with blade morphing.
Hover and forward flight tests were conducted to investigate the mutual aerodynamic interaction between the main motor and fuselage of a conventional helicopter configuration. A 0.15-scale Model 222 two-bladed teetering rotor was combined with a 0.15-scale model of the NASA Ames 40x80-foot wind tunnel 1500 horsepower test stand fairing. Configuration effects were studied by modifying the fairing to simulate a typical helicopter forebody. Separation distance between rotor and body were also investigated. Rotor and fuselage force and moment as well as pressure data are presented in graphical and tabular format. Data was taken over a range of thrust coefficients from 0.002 to 0.007. In forward flight speed ratio was varied from 0.1 to 0.3 with shaft angle varying from +4 to -12 deg. The data show that the rotors effect on the fuselage may be considerably more important to total aircraft performance than the effect of the fuselage on the rotor.
A wind tunnel program has been conducted in the NASA Ames National Full-Scale Aerodynamics Complex to investigate the aerodynamic interactions between a full-scale helicopter main rotor and simple fuselage. This program provides the first complete, quantitative data base detailing the effects of a rotor on a full-scale fuselage. Steady fuselage and rotor loads were obtained as well as steady and unsteady fuselage pressures. Data were acquired over a range of rotor thrusts, advance ratios, and shaft tilts. As expected, results indicate that the largest effects on mean fuselage forces, moments, and surface pressures occur at low speeds and high rotor thrusts. Quantitative comparisons were made to determine the ability of small-scale tests to simulate the effect of the rotor on mean fuselage forces and pressures.
Operating and aerodynamic characteristics of rotor in autorotative flight analyzed for reentry from orbit
Interactional aerodynamics of multi-rotor flows has been studied for a quadcopter representing a generic quad tilt-rotor aircraft in hover. The objective of the present study is to investigate the effects of the separation distances between rotors, and also fuselage and wings on the performance and efficiency of multirotor systems. Three-dimensional unsteady Navier-Stokes equations are solved using a spatially 5th order accurate scheme, dual-time stepping, and the Detached Eddy Simulation turbulence model. The results show that the separation distances as well as the wings have significant effects on the vertical forces of quadroror systems in hover. Understanding interactions in multi-rotor flows would help improve the design of next generation multi-rotor drones.
Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gauge output power spectra. The combined damping consists of aerodynamic and structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given equivalent speed, equivalent mass flow, and pressure ratio while structural and mechanical damping are assumed to be constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third stage rotor blade aerodynamic damping is presented and discussed for 70, 80, 90, and 100 percent design equivalent speed. The compressor overall performance and experimental Campbell diagrams for the third stage rotor blade row are also presented.
An enhanced integrated aerodynamic load/dynamic optimization procedure is developed to minimize vibratory root shears and moments. The optimization is formulated with 4/rev vertical and 3/rev inplane shears at the blade root as objective functions and constraints, and 4/rev lagging moment. Constraints are also imposed on blade natural frequencies, weight, autorotational inertia, centrifugal stress, and rotor thrust. The 'Global Criteria Approach' is used for formulating the multi-objective optimization. Design variables include spanwise distributions of bending stiffnesses, torsional stiffness, nonstructural mass, chord, radius of gyration, and blade taper ratio. The program CAMRAD is coupled with an optimizer, which consists of the program CONMIN and an approximate analysis, to obtain optimum designs. The optimization procedure is applied to an advanced rotor as a reference design. Optimum blade designs, obtained with and without a constraint on the rotor thrust, are presented and are compared to the reference blade. Substantial reductions are obtained in the vibratory root forces and moments. As a byproduct, improvements are also found in some performance parameters, such as total power required, which were not considered during optimization.