By Ray Prouty | December 8, 2017
John Q. Pilot is scooting along in his helicopter doing contour flying at high speed. The visibility is not too good. It is only good enough to SUDDENLY SEE THE HIGH-VOLTAGE LINES JUST IN FRONT!
To avoid being responsible for a region-wide power failure while producing a shower of sparks that would do neither him nor his machine any good, Johnny must either pull up to go over, push down to go under or make a sharp U-turn.
Now, in TV fashion, let’s freeze-frame Johnny with a startled look on his face. Fortunately, we can make time stand still while we calmly discuss what the designers have or have not done to help him.
The design of any vehicle is an exercise in the air of rational compromise and that is nowhere more evident than in the development of a helicopter. The two primary design influences are the desire to keep the weight down so it can hover and to keep the cost down so it can compete in the marketplace. This last consideration is very important in the free world where competition exists both in the civil and the military markets.
As a result, the helicopter capabilities Johnny can call on are limited, since they were established during the early stages of design when the engineers were under constant pressure to hold down both weight and cost.
To go over those power lines, Johnny wants to use maximum rotor thrust to pull up. Despite what is generally thought, thrust is not limited by the onset of retreating blade stall. A model rotor in a wind tunnel will go gracefully into deep retreating blade stall, producing higher and higher thrust while maintaining the ability to be trimmed out with cyclic pitch. As far as I know, no one has actually gotten the thrust to peak on a model rotor.
This apparent ignoring of retreating blade stall is because most of the thrust is being generated by the blades over the nose and tail. And even after an airfoil stalls, it will still develop a high lift coefficient. The most thrust will occur on a rotor that has a high nose-up disc angle of attack, and enough collective pitch to make even the blades over the nose and tail stall. In this case, both the lift and the drag acting at a blade element will be oriented upward, adding to the total rotor thrust.
A wind-tunnel rotor in this situation should be able to generate a thrust/solidity coefficient of at least 0.3. Since a normal value might be 0.06 for level flight at sea level, this rotor could have the capability to develop a load factor of 5 Gs in flight! But will it?
No, not really. Unlike a wind-tunnel model that is held fixed in space, the helicopter with rotor thrust greater than the gross weight will follow a cured flight path as it climbs and will never reach the high nose-up angle of attack required for the 5-G pullup. Flight tests show that a maximum value of thrust/solidity coefficient of about 0.2 has been achieved, corresponding to about 3 Gs at sea level and, of course, lower at altitude. The highest values recorded were during flares from autorotation where the disc angle of attack was already nose-up.
IF higher load-factor capabilities are required, the designers will have to choose a lower value of thrust-solidity coefficient in level flight by increasing bade area — or add a wing. Unfortunately, both will hurt the hover performance but that is just one of the compromises that must be considered in the design process.
The pilot of an airplane might be concerned with breaking the wings off during a high-G maneuver. The pilot of a helicopter need not worry. As the rotor cones up under load, centrifugal forces try to bend the blades down as much as lift forces try to bend them up. Thus, steady blade bending moments are relatively low. Since most of the rest of the airframe is designed to be stout by other considerations, such as landing loads, nothing breaks unless it has already been weakened by previous fatigue damage — a subject we will come to later.
The statement that “nothing breaks” during a high maneuver might seem at odds with the V-N Diagram (Figure 23-2) used to guide the structural design and which, at least indirectly, implies that it is dangerous to fly outside of it. However, a good indication that the V-N diagram has never been considered a serious operational limit is the fact that only on recent combat helicopters, such as the U.S. Army’s Boeing AH-64A Apache, have the designers bothered to give the pilot a G-meter.
Lowering the collective and pushing over is another possible escape maneuver but it generates a control problem. If the pullup maneuver was done with 3 Gs, a mirror-image pushover would require minus 1 G; that is, both require a 2-G change rom 1 G, level flight.
Depending on the rotor design, pitch and roll control will be lost at some low load factor. For a teetering rotor, the critical point is when it is developing no thrust in zero-G flight — but even an offset flapping-hinge rotor or a hingeless rotor will lose control at some level of negative thrust.
At minus 1 G, a rotor needs a hinge offset — or equivalent hinge offset of at least 7% of the rotor radius to maintain half the control power it had in level flight. If Johnny’s helicopter does not have this much, the pushover may not be such a good idea.
A sharp U-turn requires a combination of capabilities. First, high roll acceleration and high roll rate are needed to get into the turn — and then high thrust and high power are needed to stay in it.
Maximum roll acceleration and maximum roll rate come from two different design decisions. The type of hub determines the acceleration capability. A stiff, hingeless rotor gives the highest and a teetering rotor the lowest. An articulated rotor with offset hinges is somewhere in between.
When it comes to roll rate, however, the type of rotor does not matter. Maximum roll rate is governed almost entirely by the amount of lateral cyclic-pitch travel the designers have built into the control system, and the rotor’s Lock number, which is the ratio of its aerodynamic lifting capability to its inertia.
For the same lateral-cyclic pitch, a light rotor with lots of blade area will roll faster than a heavy rotor with a small blade area. So you can see that Johnny’s ability to quickly get into the turn has been determined by early design decisions that probably had nothing to do with the immediate situation.
Once banked over, all these roll considerations become moot Now the question becomes: “Can the rotor/engine combination produce the thrust/power required to maintain this high-G turn?”
The type of turn for which the answer is most likely to be “No” is one in which the pilot tries to maintain both the initial altitude and speed. For the pullup maneuver, he could get some benefit from increased disc angle of attack to produce rotor thrust. But in a constant-speed-and-altitude turn, the disc has to be tilted nosedown to overcome the drag and the increased thrust must come form collective pitch alone As a rule of thumb, the maximum rotor thrust in this condition would be about 70% of what it would be in a pullup.
The other limit to this maneuver is the maximum power that can be delivered to the rotor. Most modern turbine engines have an automatic limit to fuel flow when the turbine-inlet temperature reaches a preset value. Thus, on cold days, the engine can put out more power than on hot days. In an emergency, the pilot would want to use the engine power up to this “topping” limit but in so doing, he might be violating the transmission limit.
On most modern helicopters, the transmission has not been qualified to accept all of the power the engine(s) can put out on a cold day at sea level Instead, in the interest of saving weight, the transmission is “derated” by designing it to accept only the power the engine(s) can produce at some high altitude and temperature, such as the 4,000-foot, 95 F (35 C) condition used by the U.S. Army.
The engine-torque gauge in a single-engine helicopter is marked at 100% at the transmission limit. On a twin-engine helicopter, 100% on each torque gauge represents half of the transmission limit. This limit has been set after bench tests in which the transmission was run at even higher power levels so there is no immediate danger of the transmission failing if asked to transmit more than 100% torque for short periods of time. However, some progressive damage might occur due to gear-tooth and bearing wear at the high temperatures generated under these extreme conditions.
At any rate, this is only a minor concern in an emergency situation. Of course, both the rotor thrust and the engine/transmission requirements are relieved if the helicopter is permitted to slow up or to descend during the turn.
One way to make the turn faster is to generate a high sideslip — that is, to make an uncoordinated turn. This introduces the possibility of exceeding the designer’s sideslip envelope that was used to design the strength of the tail boom and the fastening of side windows and doors that are subject to suction loads during high speed sideslips.
If the designer has done his job right, the sideslip envelope cannot be exceeded in flight. This is not always true. But since the designer does not give the pilot a sideslip indicator (except in flight-test aircraft) it is a limit meant to be ignored.
Even though Johnny won’t break anything in his maneuvering, he may be doing some fatigue damage as he develops oscillating loads higher than the “endurance limit” in the rotating components. This has to be a minor consideration in an emergency but it might become a major consideration sometime later if a part were to develop a fatigue crash before its scheduled replacement time.
This problem can be traced back to the original spectrum of flight conditions used in design. As part of the “design criteria,” somebody guessed at the time a typical helicopter would spend doing hover, cruise, dash and other things, such as jumping over power lines. This was used by the designers to determine how strong (and heavy) each part must be to last the required operations hours: 4,5000 for the Apache, for instance.
Let’s say someone guessed that an extreme maneuver capable of doing some fatigue damage would occur every three hours, and in actual operations, the aircraft is flown such that this maneuver, or an equivalent one, is done every hour. Then, we shouldn’t be surprised that parts develop fatigue cracks before their calculated time. This situation especially applies to aircraft being used to train for emergency procedures.
To avoid being surprised by premature failures in the field, aircraft manufacturers may have to adopt a practice already used by the engine manufacturers — a history recorder. For engines, these automatically record the time spent over a certain turbine-inlet temperature. By correlating this with test-cell results, a mechanic can determine when the engine should be pulled for overhaul.
The ultimate procedure for the airframe would be to strain gauge each fatigue-critical component and record each cycle of load above the endurance limit and by how much it was exceeded. This is a method now used during flight test in a “flight-load survey”, and allows the structures engineer to predict when each part will suffer enough fatigue damage to warrant its replacement.
This would be quite a complicated solution in operation, so it is logical that so simpler method will be used. This could include recording load factor, or extreme vibration levels or whatever other parameter is indicative of high fatigue loads.
Did John Q. Pilot avoid running into the transmission lines? Since he was a good pilot and was flying a good helicopter, sure he did — this time.