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Aviation History
1962
1962 - 0323.PDF
— FLIGHT International, I March 1962 02 -CL- 475 FLIGHT DATA 20 40 60 80 FORWARD SPEED (m.p.h.) 100 Fig 9a (above) Overall vertical acceleration due to in-flight vibration measured over the whole range of CL-475 speeds; Fig 9b (below), accelera tion and velocity due to cabin vibration at low and high speeds 325 All full-scale flight data so far have been obtained with the CL-475. Powered by a 140 h.p. Lycoming piston engine, this pioneer helicopter seats two side-by-side and accommodates a considerable amount of instrumentation. Further data are on the opposite page o s o z H) 0 10 0 /^A,^ ACCELERATIONS Ij-I BY FREQUENCIES CABIN VERTICAL. VELOCITY IP 2P 3P "*—^/~~ V-20 MPH 033 055 020 -V-94MPH 045 10 070 # ONE ROTOR REVOLUTION CABIN LATERAL VELOCITY V=20MPH 045 070 10 '-'k/VvXV-'MMPH 045 12 030 inputs. Similar data can also be plotted in a form (Fig 6) for comparison with military specification requirements. Although no attempt was made to meet military specifications, it is significant to note that by simple changes in control gradients values of response exceeding military requirements were obtained. Steady-state roll rates were also established, and plotted as shown in Fig 7. Maxi mum rates of 45°/sec can be achieved at a stick travel of 7in. One of the serious disadvantages long held against the rigid rotor is that of supposed high stress-levels. These arguments are fre quently stated intuitively, and from considerations of absolute rigidity rather than a sound analysis of an actual system. High stresses are thought to result from large cyclic bending moments in the blades introduced by body moments; what is not recognized is that the resulting bending moments in cantilever blades are not significantly different from other types when all factors are con sidered. For example, chordwise bending stiffness is not new, nor is vertical bending stiffness in the collective, or coning, sense, and the requirement for blades to sustain bending loads for these condi tions has not produced any unusual problems. This would suggest that stiffness in the cyclic sense is entirely reasonable also. Further more, it can be shown that significant reductions in coriolis effects and second-mode bending loads result when the blades are canti- levered. In the final analysis, the structure required to provide low stress-levels in a rigid rotor is no more of a design problem than in conventional helicopters. To substantiate these principles, stress measurements were made on the CL-475 by strain-gauging the blade and hub (Fig 8). Vertical bending moments were measured on the blade at four stations, and both vertical and in-plane bending measurements were taken on the hub. Considerable information was obtained under many flight conditions, including effect of e.g. movement, load factors to 1.64g, hillside operations and autorotation flares. Peak cyclic-stress values of 1,500 lb/sq in were recorded for the blades under level flight con ditions, and increased only 250 lb/sq in for each inch of e.g. offset. Manoeuvring and autorotation flares caused maximum cyclic stress values of 1,750 lb/sq in. Maximum stresses were obtained during hillside landings on 10 slopes; but these are transient levels, and present no problem relative to blade life. Measured results for the aluminium blade indicate a considerable margin below the allowable level for anticipated flight time at each stress. For the steel hub (Fig 8c) substantial margins were again found between actual and allowable stress levels. Vibration data were obtained at several locations. The overall vibration level is low and does not vary appreciably with speed up to the maximum obtained (Fig 9a). Oscillograph traces obtained from velocity transducers in the cabin area registered both vertical and lateral vibrations, and traces are shown (Fig9b)at a low velocity typical of the transition region and at maximum level speed. A frequency spectral analysis was made and the data converted to cabin accelerations (g). Accelerations due to IP vibrations are quite low, and arise primarily from rotor unbalance or maltrack. Accelerations at 2P are low at transition velocities but increase to a peak at maximum forward speed. A theoretical analysis of a three- blade rotor indicates that 2P vibrations can arise only from mechanical effects; on the CL-475 there were two sources of such vibration, both of which can be eliminated in future designs. The maximum 3P vibration of 0.1 g occurs in the transition region, where it is primarily an aerodynamic input to the blades; this component is not predominant at other velocities. Our experience to date has convinced us that significant reductions in the sources of vibrations are inherent in the rigid rotor, making possible the elimination of various attenuating devices in present helicopters. The question is often asked whether or not the rigid rotor is limited to small machines. This arises from the intuitive conclusion that increased rotor diameter will result in high blade stresses and lower blade frequencies, leading to either heavy rotor structure or increased vibration levels. Our analysis* of this question S'IOWS that certain basic premises follow a proposed increase in size. For example, it is logical to assume that disc loading increases linearly, and that tip Mach number remains constant. Rotational frequencies therefore decrease, and a constant effective dynamic pressure for the rotor blades results. To maintain a constant blade lift coefficient the blade chord must increase as the square of the scale factor. This results in lower blade aspect ratio and an increase in blade fre quencies relative to the rotational frequency. At some point, as size increases, the designer is required to in crease the number of blades, to keep blade aspect ratio—and hence, blade frequencies—within desirable limits. In other words, the tendency is for the blade relative stiffness to increase rather than decrease with vehicle size, and the designer is therefore given more latitude in his choice of blade aspect ratio and stiffness distribution. Similar reasoning for stress and rotor weight fraction will show that these factors are independent of vehicle size, allowing a relatively wide choice to the designer in selecting blade design. The con clusion that can be drawn from an analysis of vehicle size effects is that the rigid rotor is not a limiting factor in size of vehicle. Work is now progressing on a research-vehicle programme utilizing the Lockheed rigid-rotor concept, which will extend the flight envelope considerably beyond that obtained with the CL-475. Its initial phase calls for a full-scale vehicle to be tested in the NASA Ames 40ft x 80ft wind tunnel. This test is significant in that a five-degree-of-freedom mounting system will permit extensive evaluation of the Lockheed rotor at speeds up to 200 m.p.h. This research programme also involves fabrication and flight test ing of two CL-595 vehicles to demonstrate the high performance capability of this concept. These vehicles will be turbine powered and will be completely instrumented to obtain detail test data. Flight dates are scheduled for the autumn of this year. A mockup of the CL-595 has already been constructed (heading illustration). We believe the concept developed and flight-tested at Lockheed has achieved all of the objectives described earlier, and has the potential of achieving maximum utility for the helicopter. Addi tional promise is shown relative to high-speed capability, and additional studies and tests are under way to exploit this potential. We are confident that full development of the Lockheed rigid-rotor c incept will pave the way for large scale acceptance of the helicopter as a means of transportation. * /. H. Culver and J. F. Johnston, "Note on Size Effects on the Lockheed Rigid Rotor Helicopter," Lockheed Aircraft Corporation Report No 15541.
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