Advanced Rotor Airfoil Design

Several research programs have been undertaken with a view of improving heli­copter performance by careful design and optimization of rotor airfoils. With appropriate design of both the airfoil sections and the blade geometry itself, conventional helicopters can now operate at flight speeds approaching 200 kts (370 km/h) without the rotor being limited by significant stall or compressibility effects. To achieve these speeds by expansion of the flight boundary, it is necessary to vary the airfoil section along the blade to give optimal performance for the extreme in the operating regimes encountered at that blade station.

Programs of airfoil development have been conducted by many of the major helicopter manufacturers and government research organizations. As previously discussed, the NACA 0012 airfoil represents a good compromise between high maximum lift, low pitching mo­ment, and high drag divergence Mach number performance. As shown in Fig. 7.44, reducing the thickness-to-chord ratio of the airfoil gives a marked improvement to the high Mach number performance through a reduction in the drag divergence Mach number. This allows a higher forward flight speed to be obtained with the rotor prior to the onset of increased rotor torque demands. Alternatively, for a given operational forward speed the rotor tip speed can be increased without incurring penalties of drag divergence and flow separation, and so rotor solidity could be reduced, thereby saving weight. However, reducing blade thickness limits the maximum lift capability of the airfoil at low Mach numbers, and this can adversely impact the retreating blade performance of the rotor. Therefore, generally the airfoil section thickness must be maintained to give a compromised high Mach number performance, whereas the high-lift performance is much improved by adding leading edge camber. As shown previously in Figs. 7.33 and 7.34, the maximum lift capability of an airfoil improves rapidly with the addition of some nose camber, although at the expense of some modest increase in pitching moment.

Unfortunately, the addition of camber also affects the shock strength on the airfoil lower surface at higher Mach numbers, causing a reduction in the drag divergence Mach number. Nevertheless, Perry (1987) explains how careful addition of leading edge camber can restore the C/max performance back to the levels of at least the NACA 0012 when operated at the same Mach number, whilst still retaining the higher drag divergence Mach number. The high-lift capability can only be improved further by using camber more toward the rear of the airfoil, but as shown previously in Fig. 7.25, this is at the expense of a more significant increase in pitching moment. On the retreating side of the rotor disk, the high pitching moments that are associated with cambered airfoils can normally be tolerated as the dynamic pressure is relatively low. Yet, on the advancing side, the dynamic pressure is larger, and so the blade pitching moments can be significant enough to manifest in high control loads, and possibly result in flight envelope restrictions. The final design of the airfoil section is usually a compromise in this regard.

The evolution of the Boeing (Vertol) VR airfoil series is shown in Fig. 7.45 – see Dadone and Fukushima (1975) and Dadone (1978,1982, 1987). The VR-12 and the VR-15 airfoils represent the best compromise in terms of maximum lift capability at the lower Mach num­bers typical of the retreating blade whilst also maximizing the drag divergence Mach number and meeting hover requirements and control load limitations. These sections were designed with the aid of numerical methods using a potential flow/boundary layer interaction analysis

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and a viscous transonic analysis. These analyses were previously validated against experi­mental measurements on other airfoil shapes so that they could be used with confidence in the airfoil design process.

The ONERA has conducted a systematic development of helicopter airfoil sections [see Thilbert & Gallot (1977)]. These airfoil shapes are designated as the OA – family, for which a selection is shown in Fig. 7.46. The philosophy behind the design of these airfoil shapes also follows that of high Qmax capability at low Mach numbers and a higher drag divergence Mach number. The OA-206 is an example of a thin supercritical-like section, which exhibits a higher drag divergence Mach number and gives potentially large improvements in advancing blade performance. The OA-209 is an example of an airfoil that is a compromise between advancing and retreating blade requirements, with gains in C/max relative to the NACA 0012 at low Mach numbers and with some modest increase in the drag divergence Mach number. Recall that a high Qmax capability is required only on the outboard sections of the

Diaue vuciwcen jj ana ouvo ш rauius; ana maximum lining penormance is reiaxea іиішег inboard. Thus, some reflex camber can be used on the inboard sections to generate lower pitching moments, that is, the OA-212 and OA-213 airfoils.

The Royal Aircraft Establishment and Westland Helicopters have conducted a systematic development of helicopter sections since the late 1960s. A good review of this work is given by Wilby (1979,1998). The first airfoil in the series, the RAE (NPL) 9615, used nose camber to give a moderate increase in C/max compared to the NACA 0012, with a small increase in drag divergence Mach number. Both improvements were made with only a small increase in pitching moment. Later airfoils that were developed included the 12% thick RAE 9645 (see Fig. 6.22), which has more aft camber and gives a 30% increase in Qmax relative to the NACA 0012. The RAE 9648 is a 12% thick reflexed airfoil, which gives a significant nose-up pitching moment whilst retaining most of the high-lift advantages of the RAE 9645. The RAE 9634 airfoil is a thinner 8% thick section, which is designed to minimize transonic

OA-209

flow effects by giving a higher drag divergence Mach number and to delay the nose-down pitching moment (Mach tuck) trend to as high a Mach number as possible.

A series of high-lift low pitching moment airfoils have been devised by the US Army and NASA for helicopter applications. These are designated as the NASA RC-series – see Bingham (1975) and Bingham & Noonan (1982). The RC(3) airfoil families use a careful combination of nose camber, trailing edge reflex camber and a supercritical type thickness distribution to extract the highest static C/max from the airfoil whilst retaining a very low pitching moment and a high drag divergence Mach number. The NASA RC(4) and RC(5) series were designed by Noonan (1990) for high maximum lift coefficients and are suitable for the inboard part of the blade – see Fig. 6.17. The RC(5) family has a lower thickness than the RC(4) family forward of the point of maximum thickness. The RC(6) series is described by Noonan (1991) and is a development of the RC(3) series designed for application at the tip of the blade.

It is in the more precise design optimization of airfoils to meet the 3-D unsteady flow environment at the rotor where future research challenges lie in airfoil design. This will see benefits from advanced computational fluid dynamic (CFD) models based on the Navier – Stokes equations – see Chapter 14. Design for maximum lift and low drag is important for all helicopter airfoils, but CFD methods have not yet matured to a level where turbulent flow separation and stall effects can be predicted with acceptable accuracy. The future, however, will see helicopter airfoils designed more specifically to meet the highly unsteady, 3-D flow requirements in which they really operate. This may provide an exciting opportunity to finally realize desired airfoil performance levels (Fig. 7.47) and so to produce significant gains in rotor efficiency and overall helicopter performance. Until then, the extreme operat­ing conditions and often highly unsteady flow environment found on helicopters means that rotor airfoils must be tested in a wind tunnel to fully assess both their steady and unsteady

Figure 7.47 The design of new helicopter airfoils will continue to require the synergistic use of modeling and experiment to realize desired performance gains.

aerodynamic performance. Wind tunnel tests are very expensive and time consuming, but they will always form an essential part of validating CFD models used for airfoil design.

The use of passive or active flow control devices may offer further gains in airfoil (and rotor) performance, perhaps increasing the rotor FM by 5-10% and expanding the forward flight and maneuver boundaries of the helicopter. One tried approach is to design the airfoils for natural laminar flow (NLF) over a significant portion of the leading edge region by favorably altering the pressure gradients, in principle lowering skin friction drag at lower lift coefficients. The resulting shapes, however, are usually ill-suited for helicopter rotors because they give a lower critical Mach number and/or higher pitching moments. Anyway, the natural erosion of the leading edge of the blade during service also tends to promote an earlier transition to a turbulent boundary layer, so any gains with NLF are still lost. Laminar flow on better suited rotor airfoils may be increased by using vortex generators [see Kerho & Kramer (2003)] or by applying boundary layer suction or heating/cooling to the surface. The idea here is to suppress the natural growth of unstable disturbances in the laminar boundary layer that eventually lead to transition. While demonstrated on fixed – wings [see Joslin (1998)], suction has not yet been used on helicopter rotors because of the weight and complexity of using pumps, valving, tubing, and so on. Receptivity control, which attempts to cancel out natural disturbances in the boundary layer, also offers some promise in increasing the extent of laminar flow on the blades. Such an approach, however, has many practical challenges in its implementation [see Saric et al. (1998)] and especially so for helicopter rotors. The use of zero-mass flow synthetic jets [see Amitay et al. (1998) and Hassan et al. (2002)] may also lead to enhanced rotor performance by increasing airfoil maximum lift and decreasing drag. There are also good possibilities for using these devices for airframe drag reduction. In general, active flow concepts are likely to see much future interest in the quest for better helicopter performance.