A Comparison of Interactional Aerodynamics Methods for a Helicopter in Low Speed Flight

Recent advances in computing subsonic flow have been applied to helicopter configurations with various degrees of success. This paper is a comparison of two specific methods applied to a particularly challenging regime of helicopter flight, very low speeds, where the interaction of the rotor wake and the fuselage are most significant. Comparisons are made between different methods of predicting the interactional aerodynamics associated with a simple generic helicopter configuration. These comparisons are made using fuselage pressure data from a Machscaled powered model helicopter with a rotor diameter of approximately 3 meters. The data shown are for an advance ratio of 0.05 with a thrust coefficient of 0.0066. The results of this comparison show that in this type of complex flow both analytical techniques have regions where they are more accurate in matching the experimental data. Introduction Rotorcraft configurations have always presented a challenge to the accurate prediction of vehicle aerodynamic performance. Complex lifting surfaces operating in a characteristically unsteady environment present challenge enough, but coupling this flow with the shapes common to helicopter fuselages amplifies the complexity. Often the prediction of isolated rotor aerodynamics is coupled using superposition with linear aerodynamic fuselage analyses or measured isolated fuselage data. Previous studies (reference 1 and 2) of complete helicopter configurations have shown weaknesses in the linear superposition assumptions commonly used in the design cycle. However, accurate models for the complex aerodynamic interactions between the rotor and the fuselage have not been developed as general tools available to the helicopter designer. Non-linear interaction effects arise in the aerodynamics of helicopter configurations in several cases. Among these cases clearly the wake of the rotor affects the fuselage onset flow. The wake geometry in most inflow models is assumed to be undisturbed by the fuselage. The wake does distort due to the presence of the fuselage. This distortion increases as the wake skew angle decreases (at lower speeds where the wake washes over the body). The influence of the fuselage on the inflow to the rotor is also potentially significant. The additional inflow distortion to the presence of the fuselage produces a change in the aerodynamics of the blades. This effect changes the strength of the shed wake and contributes to additional distortion of the wake. Notation Ct Thrust Coefficient, thrust/(density x disk area x tip speed x tip speed) Cp Pressure Coefficient, (pressure freestream pressure)/freestream dynamic pressure R Radius of the rotor, 1.574 m (62 in.) X Downstream length, m (in) Y Lateral distance, m (in) Z Vertical length, m (in) A1s Longitudinal Flapping angle, degrees Experimental Data The Langley 14by 22-Foot Subsonic Tunnel The Langley 14by 22-Foot Subsonic Tunnel is a closed-circuit, single-return, atmospheric wind tunnel (figure 1). In 1970 the unusual test requirements associated with V/STOL and rotorcraft aerodynamic research led to design and construction of this tunnel. The tunnel has a test 33.2 section that can be operated in a variety of configurations: closed, slotted, partially open, and open. The closed test section is 14.5 ft high by 21.75 ft wide by 50 ft long with a maximum speed of about 338 feet per second (fps). The open test section configuration, which has a maximum speed of about 270 fps, is formed by raising the ceiling and walls to form a floor-only configuration. The tunnel may be configured with a moving-belt ground plane and a floor boundary-layer removal system at the entrance to the test section for ground-effects testing. During this investigation the tunnel was configured with the walls and ceiling in the raised position to improve the quality of the low speed flow. Figure 1: Langley 14by 22-Foot Subsonic Tunnel The tunnel is equipped with an on-line static data reduction system that can display computed aerodynamic coefficients with interactions and wall interference corrections in real time. The tunnel has support, drive, and instrumentation to facilitate powered rotorcraft testing and has been used for rotorcraft investigations since its inception. In addition, the tunnel has flowvisualization and acoustic testing capabilities. Rotor Test System The rotor test system used for the experimental data reported here is built on a generic test system developed at the 14by 22-Foot Subsonic Tunnel. This test system, the General Rotor Model System (GRMS), consists of two synchronous electric motors, a combining gearbox, collective and cyclic blade pitch controls and a four-bladed articulated hub mounted on a six-component strain-gauge balance. It also includes a fuselage skin mounted on a separate, similar balance. These two six-component straingauge balances are used to provide independent measurement of the rotor and fuselage aerodynamic loads. The system as tested is shown in figure 2 installed in the wind tunnel. Figure 2: GRMS and Fuselage in Langley 14by 22-Foot Subsonic Tunnel The rotor system tested consisted of four rectangular blades on an articulated hub. The blades have a linear twist of -8.0 degrees from the center of rotation to the tip. The chord of 4.25 inches and radius of 62.0 inches gives the system a solidity of 0.087. The shape of the fuselage is designed to be representative of a wide range of helicopter fuselages without being specific to any one. The fuselage can be described by a set of superellipse equations that simplifies development of computer models. The geometry of this fuselage is described in references 2 and 3 and is shown in figure 3.