By Ray Prouty | November 3, 2017
“The pitch-link loads are too high!” This dreaded statement — prompted by looking at the data from an instrumented helicopter’s early test flight — is all too common.
It is common because we still do not know where all these loads come from. Therefore, the designer often does not receive the needed guidance to make the pitch links and the other rotor control-system components strong enough to indefinitely resist the fatigue loads that may be generated in flight.
The pitch link is the vertical — or nearly vertical — member that connects the blade pitch horn to the rotating swashplate (Figure 50-1). Its sole function is to make sure the blade responds to the collective and cyclic control commands corresponding to the swashplate’s position.
It is generally a simple rod with self-aligning bearings at each end. This means that the pitch link carries only tension and compression loads, since bending moments must be zero at each end. This simple structure makes it easy to equip with strain gauges, and interpreting the results as loads in the control system is relatively straightforward.
How does the pitch link get loaded up high enough to surprise even experienced helicopter engineers? We know some answers, but as yet not all of them. In the first place, we know of several apparent sources that turn out to be relatively innocent.
You might think that because the blade has inertial about its feathering axis, it would mean a resisting load would be produced as the blade itches up and down each revolution. However, a study of the dynamics, including the role of centrifugal forces acing on each blade element, shows that the cyclic feathering motion with a frequency of once per revolution actually puts the blade into resonance.
A system in resonance theoretically requires no force to keep it swinging. Therefore, the oscillating pitch-link load due to cycle feathering should be zero, although there will always be a steady force proportional to the collective pitch that is trying to force the blade toward its flat-pitch position.
The absence of an oscillating pitch-link load, due to the blade being in resonance, would only strictly be true if the rotor were operating in a vacuum. Going through cyclic pitch in air does introduce aerodynamic damping, due to the rate of pitch change. But this can be shown to be relatively small for a blade pitching through only one cycle during each revolution.
Another possible aerodynamic source is the change in the airfoil’s aerodynamic pitching moment as the angle of attack changes. But most designers of helicopter blades are careful to select airfoils that have little or no pitching moment about the airfoil’s aerodynamic center — the point where the lift can be assumed to be concentrated and which is on or near its quarter chord — in the normal ranges of angle of attack and Mach number where the rotor operates.
For many years, the preferred airfoils for autogyro and helicopter blades were symmetrical, such as the NACA 0012 and 0015, because they have airfoils to take advantage of their higher lift characteristics. They were still careful to modify them by bending the trailing edge up (reflexing) to nullify their basic pitching moment.
Additionally, the designer lines up both the center of gravity of the blade and the feathering axis on the quarter chord so that neither flapping motion nor changes in lift are expected to produce pitching moments on the blade. Thus, once again, web have failed to identify a significant source of high pitch-link loads and the designer might feel justified in saving eight in this component.
Note that the above discussion was limited to “normal ranges of angle of attach and Mach number.” Outside these limits, any airfoil will produce aerodynamic pitching moments either due to stall or due to compressibility effects. In each case, the result is a nosedown pitching moment as the low-pressure area on the top of the airfoil moves toward the trailing edge. Thus, at high speed we would expect to generate nosedown aerodynamic pitching moments on the advancing tip due to compressibility, and on the retreating tip due to stall.
It would be nice to be able to say that the aerodynamicist could predict these accurately enough to be able to give the designer realistic loads for all the elements of the control system, including the pitch links. Unfortunately, the complicated aerodynamic in both of these regions are still not well understood.
One result of this lack of understanding can be seen on the Boeing AH-64 Apache. As first flown, the blades had straight tips. At high speeds, the loads at the hydraulic-control actuators were significantly higher than had been designed for. The problem was traced to a nosedown “spike,” generated as each blade encountered compressibility on the advancing side.
Rather than redesign the actuators that had just gone through a long and expensive development and qualification process, it was decided to change the blade design by incorporating sweep to fool the tip into thinking it was at a lower Mach number. This solved the problem, though at the expense of higher blade cost.
If the retreating tip is in deep stall, as it might be in high-speed level flight or when pulling high load factors at any speed, the resulting nosedown pitching moments will also generate loads in the control system.
A typical result of flight tests on a new helicopter is shown in Figure 21-2, where cyclic load sin the pitch link are plotted against forward speed. At some speed, the loads, which in this case are attributed to retreating blade stall, are above the endurance limit. This means that they would eventually start fatigue cracks if much flying were done above the critical speed.
There is another source of potential stall, which primarily happens during descents or flares, when one blade flies through, or close to, the trialing tip vortex shed by the blades ahead of it. The resulting high angles of attack can produce stall over a local area with resulting pitching moments. This effect adds to the “rocky road” roughness that most helicopters encounter during a landing flare.
If the blade is attached to the hub through a strap pack as on most McDonnell Douglas helicopters, or with a flexbeam as on many of the new bearingless rotors, feathering motion will produce control-system loads as these components are twisted. To minimize the steady force, it is fairly common to install the blades such that the strap pack or flexbeam is not twisted at a collective-pitch value corresponding to cruise flight.
All of the above discussion has ignored the fact that the blade may not remain a straight piece of structure while being subjected to many complex aerodynamic and dynamic effects. If you have ever seen movies of a blade taken from a hub-mounted camera, you know that it does a wild dance out there.
As a matter of fact, a normal long and limber blade has more characteristics of a chain than a beam, since most of the stiffening is produced by centrifugal forces rather than by structure. A chain rotating in a plane has a natural frequency close to three times per revolution with a “mode shape,” and so do most blades.
A blade bending flapwise introduces a new source of pitch-link loads in that the drag on at least part of the blade will have a moment arm around the feathering axis to produce either compression or tension in the pitch links.
Similarly, chordwise bending offsets the aerodynamic center from the feathering axis and gives lift a change to produce a pitching moment. At the same time, accelerations associated with flapwise bending, chordwise bending and torsion will produce inertial forces that may couple in odd ways and will add to or subtract from the loads that get into the control system.
You can see that the loads engineer — who should be a combination of an aerodynamicist and a dynamicist — has a difficult job if he is o accurately predict the fatigue loads that should be planned for when designing the aircraft. As far as I know, no one has yet been able to satisfactorily calculate pitch-link loads that agree with measured pitch-link loads. For that reason, the loads engineer often will try to make a logical scaling from some previous helicopter.
Historically, this has not been a very successful approach. Parts have had to be redesigned, as on the Apache, or elaborate monitoring schemes, such as the Cruise Control Indicator on the Boeing CH-47, have had to be developed to help the pilot avoid flight conditions where the loads can do fatigue damage.
The hope for this situation may be in the exploitation of some of those mysterious dynamic couplings that sometimes increase the oscillatory loads and sometimes decrease them. If by a good understanding of these effects we can emphasize the latter while minimizing the former, we will have a better chance of not only avoiding the need to design changes after the helicopter starts its flight tests, but of actually being able to use smaller and lighter parts in the control system.