Just as a horizontal stabilizer is not absolutely necessary on a helicopter, neither is a vertical stabilizer, since the tail rotor alone can produce adequate directional stability. Nevertheless, most modern helicopters do have a vertical stabilizer. Depending on the helicopter, the vertical surface may be used to: streamline the tail rotor support, supplement the directional stability produced by the tail rotor, unload the tail rotor in forward flight, support a T-tail, or stabilize the fuselage in case the tail rotor drops off completely.

Because of the low dynamic pressure behind the fuselage and hub, as illustrated by the measured data of Figure 8.9, many designers use vertical surfaces on the ends of the horizontal stabilizer to put them into relatively clean air. In this configuration, they are also out of the way of the tail rotor’s induced flow, and

they increase the effectiveness of the horizontal stabilizer by acting as end plates. This configuration, however, is probably heavier than a centrally located fin.

As long as the vertical surface is there, it can be used to unload the tail rotor in forward flight by including camber or an offset incidence angle. The primary purpose is to increase the fatigue life of the tail rotor by minimizing the oscillatory flapping loads. Unloading the tail rotor in this manner may not save on total power, since now the induced drag of the vertical surface must be overcome by the main rotor. If the span of the vertical surface is much less than that of the tail rotor, the power tradeoff will probably be unfavorable.

Some recent specifications for combat helicopters have asked the designers to configure the aircraft so they could be flown home in case the tail rotor were completely shot off. In a sideslip, a big enough vertical surface could produce enough antitorque force to do this. Unfortunately, several development programs, such as those reported in references 10.8 and 10.13, have demonstrated that this much area can cause large blockage problems, especially in sideward flight. As a result no helicopter has this desirable characteristic at this writing.

A wing is a convenient place to hang external stores or to carry fuel. From an aerodynamic standpoint, however, it is usually detrimental unless used in conjunction with a forward propulsion system in a compound helicopter configuration. The penalty in hover is due not only to the structural weight of the wing but also to its aerodynamic download. In high-speed forward flight, a lifting wing that is used to unload the rotor may actually increase the retreating blade angle of attack, leading to premature stall. This happens because the partially unloaded rotor must be tilted further forward to produce the required propulsive force that now must compensate for the drag of the wing in addition to the drag of the basic helicopter. This requires increased collective pitch to overcome the inflow and increased cyclic pitch to keep the rotor in trim. In many applications, these two effects will combine to increase the retreating tip angle of attack more than the decrease made by unloading the rotor. In some cases, if the requirement is for a high transient load factor during a pullup, then the addition of a wing may be justified.

If a wing is to be used, its incidence should be chosen so that at high speed it is operating at its best angle of attack. By analogy with biplanes, for minimum induced drag, the wing and the rotor should be sharing the lift such that their ratio is:

The wing should be located with its aerodynamic center behind the most aft center of gravity of the helicopter so that it acts as a stabilizer rather than a destabilizer.

A large wing may cause a problem in autorotation. If the wing supports a large portion of the gross weight, the rotor will be starved for thrust and will not be able to maintain autorotation. If this situation is possible, some means of reducing wing lift will have to be used, such as reverse flaps, spoilers, or incidence changes.

Auxiliary propulsion in the form of propellers, ducted fans, or jet engines can be used to relieve the rotor of part or all of its propulsive requirement. The limit to this is the wingless autogiro, where no power at all is required by the rotor. The use of auxiliary propulsion to overcome drag reduces the needed forward tilt of the rotor, thus decreasing its collective and cyclic pitch requirements and relieving the high angle-of-attack pattern on the retreating side. This permits this type of aircraft to operate at very high tip speed ratios. Autogiros built in the 1930s routinely operated at tip speed ratios above 0.5, which is considered today to be the upper limit for "pure” helicopters.

An auxiliary propulsion system acting as a separate, controllable longitudinal force system can be used to change the aircraft pitch attitude at any speed. This is especially attractive for combat helicopters. In a hover, for example, this type of helicopter can be held either nose up or nose down to increase the usable field-of – fire. Another use is as a speed brake in forward flight. This is a valuable attribute where rapid and precise decelerations and descents are required to accomplish a mission.

The power to operate the propulsive devices depends on their thrust-to – power ratio. For propellers and ducted fans, this ratio is primarily a function of their diameter—the bigger, the better. Such a system does have its disadvantages, of course. Its weight subtracts from the payload that can be hovered; unless it is declutched in hover, it will reduce the power available to the main rotor, thus creating an even higher payload penalty.

The compound helicopter is even more vulnerable than the pure helicopter to the generality that whatever helps the high speed capability hurts the hovering capability, and vice versa.

The complexity and weight penalties of compounding by using both a wing and an auxiliary propulsion system are justified only when the aircraft require­ments combine higher speeds than can be achieved by a pure helicopter with helicopter-type performance in the lower speed regimes. If the rotor is completely unloaded, all the limitations based on blade element angles of attack are eliminated. At some forward speed, the advancing tip Mach number will approach drag divergence; at that point, however, the rotor can be slowed down to alleviate even this problem.

Despite the possibility of unloading the rotor completely, analysis shows that a lifting rotor acts as an efficient large-span wing compared to the typical wing that might be used on this type of configuration. For most effective use of the power, the main rotor should be loaded as highly as possible. This is done by adjusting the collective pitch to some value above flat pitch. In this sense, the collective control is something like the flap setting on an airplane.

The U. S. military has developed a standard weight reporting format known as a Group Weight Statement. The weight equations given earlier have been applied to the example helicopter and its Group Weight Statement is included in Appendix A.

The best position for the center of gravity on a single-rotor helicopter is slightly ahead of the main rotor shaft. Even when the designer has good intentions and achieves this goal during preliminary design, experience indicates that before the helicopter is ready for production, its average center of gravity (C. G.) will have drifted aft and settled down somewhere behind the shaft. In some cases, this aft C. G. position is forced by design considerations even during preliminary design. A classic example of this is the Sikorsky UH-60A. As explained in reference 10.13, this helicopter was limited in overall length by the requirement to load it into a C-130 without major disassembly. With the rotor sized by the vertical climb requirement, the air transportability requirement dictated a short nose. This and the desire to carry all the fuel behind the main cabin, led to a center of gravity range of more than 15 inches; all located aft of the main rotor. Sikorsky more or less satisfactorily solved the resultant trim problem by canting the tail rotor, as explained during the discussion of tail rotor design.

The longitudinal position of the center of gravity of the empty helicopter is calculated from the sum of the static moments about some arbitrary point contributed by each group that makes up the empty weight divided by that weight.

The arbitrary point should be selected ahead of and below the nose so that all parts of the aircraft will have positive locations. Figure 10.10 shows the scheme as it applies to the example helicopter. Since the center-of-gravity position with respect to the main rotor is of prime importance, it is convenient to chose the origin so that the hub falls on an easily remembered point. For the example helicopter, this has been chosen as Fuselage Station 300 and Waterline 200. The calculation that yields the center-of-gravity position for the empty helicopter is presented next.

Calculation for Center-of-Gravity Position of Empty Example Helicopter








Moment (in.-lbs)













Alighting gear












Flight controls




Auxiliary power plant




















Furnishings and equipment




Air cond & anti-ice




Manufacturing variation




Total Empty Weight



Total moment: 3,056,450

C. G. pos.^pjy —

— 305.6 fuselage station


Note that as the design progresses, this calculation can become more and more precise by expanding it to account for the weight and location of each component of each group.

A plot of the longitudinal center of gravity position as a function of gross weight is known colloquially as the "C. G. potato.” It is generated by loading items of the useful load into the empty helicopter in the most forward manner and then in the most aft manner. Figure 10.10 shows the location and weights of the useful load items of the example helicopter. The fuselage station for the center of gravity as the loading proceeds is obtained from the equation:


Total Moment, empty + ^ JF„(Fuse. Sta)„

Fuse. Sta. n = ———————————- ——^——————

Empty Weight + J Wn

n =1

The resulting C. G. diagram is shown in Figure 10.11 (page 667) for loading starting from the front and for loading starting from the rear. Any other loading sequence would place the center of gravity inside the diagram. For this case, the maximum possible C. G. excursion is 28 inches ahead and 37 inches behind the main rotor. This extreme result of "indiscriminate loading” might be marginally acceptable for the forward condition, where a nose-up rotor flapping of about 6° would be required to trim, with substantial resulting blade loads. The far aft position would certainly be unacceptable because of even greater flapping as well as the destabilizing effect of the aft C. G. In operation, a helicopter of this size would probably be limited to not more than ±10 inches from the shaft, which means that the loading would have to be monitored and controlled by the crew.


The following items can be evaluated by the methods of this chapter:

Center-of-gravity position



Component weights


Growth factor


Minimum tail rotor solidity


Optimum ratio of wing lift to rotor thrust


Ratio of useful load to gross weight


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