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Aviation History
1962
1962 - 0322.PDF
324 FLIGHT International, 1 March 1962 LOCKHEED'S HELICOPTER COLL STABILITY 20 DAMPING INERTIA L-475 FLIGHT DATA ROLL CONTROL ggg*gjgS Fig 3 Measured roll dynamic stability for the CL-47S, compared with that of an S-51 helicopter .NORMAL C.6.RANGE JNSAFE H UNSAFE SAFE HIN6ED ROTOR LOCKHEED UNSAFE* •* Sg, UNSAFE ROTOR TYPE L "«SAFE-?a Fig 4 Lockheed claim much greater flexi bility in safe loading STICK DEFLECT ION- INCHES RATE OF ROLL- DE6/SEC 2 1 0 12 8 4 0 ('. . . •A -8 12 16 TIME (s«c) 20 20 2 2 3 a 15- I" UI z - 5 ( ... • -ORIGINAL 8 PRESENT^ g MIL. SPEC. •§ * BOUNDARY 2 > 1 2 3 4 5 6 DE6/IN /'ASEC 7' 6 i; 2 ORIGINAL • PRESENT ' 5 12 3 4 5 6 DEG/IN/SEC Fig S Response of the CL-475 to a suddenly induced roll demand Fig 6 Similar response data to Fig 5 plotted in the form of stability against rate of roll 6 6- 4- 1l- V'50-60 MPH S u FLIGHT 182^ m 2" (3-25-61)^ 2 4- LIGHT 114 , "FULL.THROW , B50 40 30 20 10 0 K) 20 30 40 50 L RATE OF ROLL-DE6/SEC R Fig 7 Measured points showing the steady- state roll capability of the Lockheed CL-475 had a 64in rotor powered by an electric motor in the fuselage. The rotor and upper control system were dynamically scaled with respect to mass and stiffness distributions, to reproduce correct dynamic and aeroelastic characteristics. The blades were strain- gauged to measure both spanwise and chordwise stresses, and instrumentation was installed to measure instantaneous blade feathering angles. The data obtained further substantiated the low vibration and stress levels of the Lockheed rotor, and provided excellent basic data on its dynamic characteristics. These results encouraged Lockheed to build a full-scale flight vehicle. This was given the model number of CL-475, and was built on company funds solely to prove the basic principles of the concept and demonstrate the low vibration levels that could be achieved (a photograph appears on the opposite page). It is a small two- seater. The 140 h.p. engine drives a 32ft three-blade rotor at 300 r.p.m., the airfoil section of the blades being a NASA 0012 with a chord of 12in. Gross weight is approximately 2,0001b, and maximum level speed is power-limited to about 90 m.p.h. Strain- gauges on the blades and vibration transducers in the fuselage pro vided data on rotor stress levels and cabin vibration levels, a total of 32 channels being recorded in over lOOhr of flight time covering a wide variety of conditions. Tests were conducted on handling characteristics, stability in hovering and forward flight, autorotation characteristics including autorotation flares to landing, hill-side VTOL, response to step control motions, and effects of wide e.g. movements. During many flight demonstrations military and civilian personnel with varying degrees of piloting skill flew the CL-475 with little or no checkout time, and on several occasions people with no helicopter experience flew after less than 30min discussion on the ground. These experi ences, and the recorded data, demonstrate the capability of the Lockheed helicopter in meeting the desired objectives. Data obtained with the CL-475 are summarized in Figs 3 to 7. The first diagram plots stability and control data in the form of damping versus control power. Whereas conventional helicopters are generally limited to values shown as a heavy shaded area on this diagram, the CL-475 exhibits values an order of magnitude higher. As a matter of interest, the larger shaded area indicates the region available to the designer with only minor changes in control-gyro inertia or control force gradients. Another significant advantage of the control power available is reflected in the increase in the allow able e.g. range (Fig. 4). Indiscriminate loading is possible on a machine of this type with no danger of upset possible. Side- hill landings and landings in gusty winds are greatly simplified as a result of the attitude stability and control power of the Lockheed rotor. Oscillograph traces (Fig 5), typical of the response of the vehicle to approximately step input stick motions, are plotted for stick position and rate of roll for a left-roll manoeuvre. The significant aspect is the highly damped and rapid response to step control Fig 8a (far left) An aluminium blade instru- mented for stress measurement at four stations during CL-475 flying; Fig 8b (left), the instru mented steel hub; Fig 8c (below), measured stress levels for three flight cases for the hub shown in Fig 8b 120 90 CYCLIC STRESS 60 (Ib/sq in xlOOO) 30 ALLOWABLE FATIGUE LIFE FOR STEEL - HILLSIDE LANDING £ TAKE-OFF 10* SLOPE pl.6q CYCLIC STRESS 1.0 q CYCLIC STRESS -j 10 10 100 FLIGHT TIME (M 1000 10,000
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