Producing Maximum Load Factors
The pilot who has been asked to demonstrate the maximum load factor capability of a helicopter will build up the maneuver until he reaches one of the following limits:
• Maximum engine power
• Maximum stick displacement
• Unacceptable level of vibration
• High nose-up attitude or pitch rate from which recovery is uncertain
• Indication of abnormally high loads in the rotor or the control system
• Aircraft instability
• Ominous change in the rotor noise level
• Sudden rotor out-of-track condition
Some of these limits—such as reaching maximum engine power—are straightforward and an be predicted by methods already developed in previous chapters. Others, however, are a function of the structural and dynamic characteristics of the rotor, the control system, and the remainder of the helicopter and of the pilot’s willingness to subject himself to uncomfortable or potentially dangerous flight conditions. There is as yet no analytical method for predicting the maximum attainable load factor when these latter considerations are involved, but there are enough experimental data to provide some insight into the problem. A convenient nondi – mensional representation of the maximum thrust capability is a plot of CT/cr versus |X such as Figure 5.2. (Note: if the blade is not of constant chord, the definition of thrust-weighted solidity given in Chapter 1 should be used in this analysis.) The plot has three boundaries depending on the flight condition.
The transient boundary an be achieved momentarily in flight or continuously in a wind tunnel at high rotor angles of attack and corresponds to every blade element operating at its maximum lift coefficient. Test results indicate that this boundary is in the neighborhood of CT/o = 0.17, which is equivalent to an average lift coefficient of over 1. Values of this magnitude or higher have been
О.08 .06 .04 .02
.05 .10 .15 .20 .25 .30 .35 .40 .45 .50
Tip Speed Ratio, ц
FIGURE 5.2 Rotor Thrust Capability
reported: in flight maneuvers by references 5.1 and 5.2; in autorotative flares by reference 5.3; and in wind tunnel tests by references 5.4 and 5.5. The theoretical limit of the transient capability is a value of CT/o of about 0.3, which could be achieved on a heavily stalled rotor at a very high angle of attack, where the contribution of the drag acting on each blade element parallel to the shaft is producing most of the rotor force. Other boundaries in Figure 5.2 represent level flight and steady turns and $re primarily based on the flight experience reported in references 5.6 and 5.7. For this reason, these boundaries probably reflect some of the nonaerodynamic limits listed in the beginning of this section. These limits vary from helicopter to helicopter and determine how heavily loaded a rotor may be before it gets into stall-related troubles. For example, it is common to generate high oscillating loads in the control system when the blade goes into stall on the retreating side. The resultant change in aerodynamic pitching moment can twist the blade nose down to an unstalled condition, only to have it spring back into stall. Depending on the torsional natural frequency of the blade and control system, each blade may go through several cycles of this oscillation—sometimes called stall flutter— while passing through the retreating side. A theoretical and experimental investigation of this problem is reported in reference 5.8 in which it is shown that control loads are higher when the torsional natural frequeqfy ratios are between 5 and 12 per revolution than for frequencies on either side of this range.
Another possible source of high vibration and high loads is the excitation of those blade-bending modes for which the forcing function Increases with tip speed
ratio and rotor thrust. One example of such a mode is the second blade flapping mode, whose natural frequency is near 3/rev (it would be exactly V27T or 2.51/rev if the blade acted like a chain stiffened only be centrifugal forces). As the tip speed ratio is increased, the 3/rev content of the aerodynamic forcing function—which, incidentally, is made larger by blade twist—increases, thus exciting the second blade flapping mode to a high amplitude with correspondingly high vertical shear at the flapping hinges. This particular vibration can be especially severe for a three – bladed rotor since all three blades will respond in unison to the 3/rev aerodynamics. On the two-, four-, or five-bladed rotor, the blades will be out of phase with each other and thus will produce less vibration in the helicopter as a whole.
Other factors that affect the apparent ability of the rotor to develop high load factors as limited by vibration are:
• The vibration isolation of the rotor and transmission from the rest of the helicopter;
• The natural frequency of the fuselage;
• The location of the cockpit with respect to fuselage nodal points (points that appear to stand still)
Thus it may be seen that there can be no single simple method for predicting the maximum load factor that will subject the pilot to unacceptable levels of vibration.
An exploration of the aerodynamic limits with a wind tunnel model is reported in reference 5.9. Figure 5.3 shows the limit of test points achieved at a tip speed ratio of 0.6. These limits are well beyond those normally accepted and were not limited by aerodynamic or structural phenomena but by straightforward mechanical interferences on flapping and cyclic pitch control on that particular model. Reference 5.9 uses these test results to predict high rotor performance for future rotors. A continuation of this wind tunnel study is reported in reference
5.10, which concludes that a 225-knot helicopter is feasible. The bottom portion of Figure 5.3 summarizes the maximum rotor thrust capability as measured on the wind tunnel model during that project.  
Source: McHugh & Harris, “Have We Overlooked the Full Potential for the Conventional Rotor?" JAMS 21 -3,1976; McHugh, “What are the Lift and Propulsive Force Limits at High Speed for the Conventional Rotor?” AHS 34th Forum, 1978
• What are the boundaries of the "Deadman’s Curve”?
• What is the minimum touchdown speed?
• How does this helicopter compare with others?
These questions can be answered using the following analytical methods.