Course Outline

1. Course Outline • Review of Conceptual Design Solutions • Conceptual Design Issues for Resolution • Structural Design • Dynamics • Stability and Control • Drive System Design • Life Cycle Cost • Power Plant Selection and Installation • Secondary Power Systems • Weight and Balance • Maintainability • Reliability and Availability • Configuration and Arrangement

2. Rotorcraft Structural Design Dr. Daniel P. Schrage Rotorcraft Design Professor Georgia Tech

3. Structural Design • Definition of structural design criteria • Determination of basic loads based upon these criteria • Demonstration by preliminary structural analysis of the adequacy of the structure to withstand these loads • Design criteria are based primarily upon military specifications, FAA standards and company experience

4. Structural Design Criteria • Developed for the following conditions: • Flight and takeoff • Landing • Ground • Controls • Special loadings • Miscellaneous • These criteria, together with the design parameters and characteristics of a given model helicopter, are used to calculate basic loads

5. Loading Conditions • Practically impossible to analyze structural adequacy for every loading condition a rotorcraft might encounter • Experience has shown that there are only a few critical loading conditions • During PD these critical conditions must be identified and the applicable loads calculated • If the rotorcraft can withstand these critical loads, it will have an adequate margin of safety for all other loads normally encountered

6. Fundamental Design Parameters • Design limit flight speed, VDL is defined as a given percentage above VH • The VH is the design maximum level flight speed in forward, rearward, or sideward flight (35kts) • MIL-S-8698 defines VH in forward flight as the max speed attainable at the basic design GW in level flight using intermediate (30 min) power, or as may be limited by blade stall or compressibility effects • The VDL/ VH shall not be < 1.25 for attack, 1.20 for utility, and 1.15 for observation, training, cargo, or heavy-lift helicopters

7. Fundamental Design Parameters • Three gross weights, Wg, are significant to the PD of rotorcraft: minimum design, basic structural design, and max alternate design gross weight • Minimum DGW is the lowest Wg considered for practical flight and represents a helicopter returning from a mission with all disposable payload items expended (5% usuable fuel,min.oil, min crew (170 lbs))

8. Fundamental Design Parameters • Basic structural design gross weight, WBSD shall be WTOw/full internal fuel & full internal & external load items required for performance of the primary mission • Max alternate design gross weight shall be as prescribed in the helicopter detail specification. In any case, the Wmax Alt shall not be greater than the max GW at which the helicopter can take off from an unprepared field of 800 ft length and clear a 50 ft obstacle in not less than an additional 200 ft @ SLS

9. Fundamental Design Parameters • The load factors applicable to Wmax Alt shall be the load factors specified for the WBSD multiplied by the ratio of the WBSD to the Wmax Alt . These load factors should not be < 2.0 • The Wmax Alt shall be used for landing and ground handling conditions to the extent specified • The range of CG locations resulting from the range of gross weights shall be considered • GWs for design shall be all critical GWs between the minimum and the maximum alternate DGWs

10. Flight Envelopes and Takeoff Loading Conditions • Forces and moments acting upon the helicopter during flight can be represented by forces and moments acting along and about three mutually perpendicular body axes(See Figure 4-1) • Any any instant in time, all aero and inertia forces shall be in equilibrium, and the following conditions shall be satisfied: • Summation of forces along each of the 3 axes = 0 • Summation of moments about each of the 3 axes = 0

12. Flight Envelopes and Takeoff Loading Conditions • Rotor thrust shall be in equilibrium with the aero and inertia forces acting upon the aircraft • For the case of trimmed 1-g flight, the rotor thrust T is equal to the resultant of the W and the aero drag D, and has a line of action in opposition to the resultant FR of these forces (See Fig. 4-2)

14. Flight Envelopes and Takeoff Loading Conditions • For flight conditions in which there is no angular acceleration, inertia and gravity forces shall be distributed in the same manner as weights, with their resultant acting through the aircraft CG • Inertia and gravity forces, for convenience in analysis, are combined as the product of a normal load factor nz and the gross weight Wg Fz = nz Wg Where Fz = normal force component, lb

15. Flight Envelopes and Takeoff Loading Conditions • The equation defining the relationship between load factornz and linear acceleration az for zero pitch attitude is: nz = 1 + az/g, dimensionless where g= acceleration due to gravity, 32.2 ft/sec2 • Additional forces and moments are created during those flight conditions in which the helicopter is being maneuvered by the pilot, or in which external forces such as gusts cause the aircraft to be accelerated in either a linear or an angular fashion • These forces and moments produce linear and angular accelerations that result in balancing inertia forces and moments. The aircraft remains in equilibrium

16. Flight Envelopes and Takeoff Loading Conditions • As an example of the stated principles, fore or aft movement of the control stick changes the direction of thrust of the rotor, resulting in an unbalanced moment about the y-axis • This moment causes the aircraft to pitch about the y-axis, and the resulting angular acceleration when multiplied by inertia results in a moment counterbalancing the applied moment, as shown in Figure 4-3 • For angular accelerations, a system similar to that for linear accelerations is used, except that the force Fz acting upon any element in the aircraft is also a function of the distance x’ of that element from the CG and is Fz = w(nz + x’θ/g), lb, where θ= ang accel, rad/sec2

18. Flight Envelopes and Takeoff Loading Conditions • Basic Flight Loading Conditions • The dynamic components – including drive shafts, transmissions, rotor hubs, and blades are sized primarily from a consideration of fatigue strength • Critical flight loading conditions( helicopter) • Max Speed (design limit speed VDL) • Sym dive and pullout at VDL and 0.6 VDL(~max n) • Vertical takeoff (jump takeoff) • Rolling pullout • Yaw • Autorotational maneuvers • These must be examined at max & min rotor speeds; also min design LF, ~0.0, shall be investig

19. Flight Envelopes and Takeoff Loading Conditions • Load factor capability can be directly in terms of rotor thrust by the method which defines LF capability of a rotor in terms of the max mean lift coefficient, CLmax. The expression for LF capability nzmax is: nzmax = (CLmax/ CLt)(B3 + 1.5Bµ2N-4/3(µ3N)) X (B3 + 1.5Bµ2t-4/3(µ3t)) (ΩN/ Ωt)2 (cosαoN/ (cosαot)3 (4-5) • The formula is conservative because the max mean rotor lift coefficient is an ideal theoretical value, usually not achievable with actual rotor blades • Modifications can be made to account for factors such as blades tapering in thickness and blade root cutout areas

20. Mission Profile and Fatigue Analysis • The design shall insure that the helicopter is capable of achieving the operational design flight requirements, is safe to operate, is free from fatigue failure throughout its design life, and has adequate reliability with a minimum of maintenance • At the same time, it is important to avoid penalizing the aircraft with excessive weight so that it cannot perform its mission efficiently

21. Mission Profile and Fatigue Analysis • Fundamental for achievement is a clear definition of the missions the aircraft is to perform • Because most helicopters will be used for a number of different purposes, it is necessary to define a variety of missions that a specific helicopter will perform during its design life • An estimate of the percentage of useful life that will be spent performing each mission also must be made

22. Mission Profile and Fatigue Analysis • A large number of variable parameters enters into the determination of fatigue life • In addition to defining aircraft missions in quantitative terms, it is necessary to make estimates of a large number of factors and to take these into consideration during generation of a fatigue loading spectrum

23. Mission Profile and Fatigue Analysis • Some of the more important variables are: • Flight altitudes • Number of takeoffs and landings per mission • Aircraft weight at takeoff, during mission and at landing • Loads due to external stores and/or cargo slings • Airspeeds • CG range • Rotor speeds • Gust effects • Magnitude, number, and duration of pilot-initiated load factors during maneuvers • Sinking speed during landing • Autorotation

24. Mission Profile and Fatigue Analysis • Two typical mission profiles for an armed helicopter are shown in Figs. 4-4 & 4-5 • Time for each event and the number & duration of events occurring during the aircraft life are determined from engineering estimates based upon the class of helicopter, field experience with that class of aircraft, and estimated performance of the vehicle itself • Percentage of flight time spent in each flight condition then can be computed from the equation in Note of Table 4-1

25. Figure 4-4 A

26. Figure 4-4B

27. Figure 4-5A

28. Figure 4-5B

29. Mission Profile and Fatigue Analysis • In addition to the spectrum of flight conditions, severity distributions of load factors due to pilot-initiated maneuvers and gust encounters are required • Fig. 4-6 is representative of the severity of load factors due to maneuvers and gust during the life of an attack helicopter and a utility helicopter, while Fig. 4-7 represents such load factors for a fixed wing aircraft

32. Mission Profile and Fatigue Analysis • Data presented in Fig. 4-6 are suitable for use during PD and analysis of attack and utility helicopters, while Fig. 4-7 is suitable for all helicopter classes (not really!) • Fatigue damage due to landings and ground conditions also must be included in the computation of fatigue life of the helicopter • Final determination of fatigue life of the airframe and dynamic components can be made only after laboratory fatigue testing of specimens and determination of the magnitude of loads during the flight load survey

33. Limit Load Factors • A clear concept of limit loads is an important factor in the efficient design and safe operation of a helicopter • The vertical accelerations that establish the maneuver limit loads can be related to the maximum thrust capability of the rotor • The maximum LF is the ratio of the max possilbe thrust that can be developed by the rotor to the gross weight • The max thrust is determined by assuming the max lift coefficient at all blade sections (CT/σ)max = Tmax/ρbceR(ΩR)2 (4-6) • The lift coefficient for the BSDGW may be represented (CT/σ)design = Tdesign/ρbceR(ΩR)2 (4-7)

34. Limit Load Factors • The max possilbe LF nzmax or max g load obtainable at BSDGW then would be obtained from Eqns 4-6 and 4-7 as an alternate to Eq 4-5: nzmax = (CT/σ)max/ (CT/σ)design (4-8) • This max attainable LF can be computed for a given rotor system and vehicle combination • However, MIL-S-8698 has established LFs to be used in the design and qualification of three different classes of helicopters • They are shown in Table 4-2

35. Control of Limit Load Factors • As flight speed increases, a given rotor AOA change produces a larger thrust increment so that large LFs may be reached w/o large attitude changes • The max loads thus may be obtained by variation in rate of control application, magnitude of control movement, and airspeed • Artificial limitation of LFs may be accomplished by: • Rotor speed regulation • Dampers in controls • Force gradients in controls • Except for rotor speed regulation, these artificial methods of limitation have not been popular for helicopter design and operation

36. V-n Diagram • Typical V-n diagrams are shown in Fig. 4-8 • The LFs shown are the limit LFs for hypothetical helicopter designs • The upper limit of 3.5 g and the lower limit of – 0.5 g in the low airspeed and nomral airspeed region were established by MIL-S-8698 • The limits at high airspeed were determined by rotor blade stall and blade tip Mach limits • The LF nz shown in this diagram is: nz = cosθ+ az/g (4-9)

37. Maneuvers (Symmetrical Flight) • Various maneuvers that are classified as symmetrical flight maneuvers include the following: • Hover • Takeoff and climb • Level flight • One-g dive • Pullup

38. Rotor Speed and Power Ranges • There is a significant difference in LF consideration between power-on and power-off operations • The allowable range of power-on rotor speed is relatively small, being limited on the high side by engine limits and on the low side by helicopter rotor blade stall and associated vibration and comfort levels • Power-off operation, on the other hand, has a comparatively large allowable range of rotor speeds • The max allowable rotor speed in autorotation is considerably higher than the max speed allowable power-on • This max power-off rotor speed limit is essentially the design limit of the rotor (with some reduction for a factor of safety) • The min power-off rotor rpm may be set at the lowest level from which, as shown by flight test experience, rotor speed recovery can be accomplished in a safe length of time

39. LFs for Other Than Normal GWs • If a helicopter is operated at a weight greater or less than the normal GW, it must be assumed that a different maneuvering load margin is available • For this reason it is customary to specify higher-than-normal GW – normally called an alternate design gross weight • The LF for this configuration should not be less than 2.0 g • The configuration that involves a lighter-than-normal GW also should be considered • This creates a condition where the relationship between rotor thrust capability and actual GW results in a higher limit LF • The principal requirement for application of this higher LF would be for certain equipment items

40. Fatigue Analysis • The effect of a particular maneuver on fatigue life must be determined by inflight measurement of the stresses on the components for which the fatigue life information is needed • Certain components may accumulate damaging fgatigue cycles during high-speed flight with a relatiely modest LF increase • On the other hand; higher LFs in other flight regimes may result in no fatigue damage • It is essential, therefore, that information on both the LF and the resulting stress is known • In addition, it is important to have an accurate measurement or estimate of the maneuver time spectrum • A typical manu time history showing LF versus time is illustrated in Fig 4-12

41. Fatigue Analysis • This type of maneuver time history may be used to determine the approx number of load cycles at each LF for the particular flight condition • The number of such maneuvers in a given time can be estimated from data • A total cycle count for LFs and associated stresses then can be accumulated to establish the effect of the maneuver upon the fatigue life of a component • This procedure would be repeated for all other maneuvers that have been determined to be critical for fatigue life of components • The PD task should include definition of the possible maneuver spectrum for the helicopter being designed • It is not possible to make a definitive fatigue analysis w/o flight test data from the actual helicopter, but background data may be used in the preliminary analysis

42. Static Analysis • The static analysis of the helicopter should be based upon one or several critical maneuvers that result in the max possible limit LFs • These maneuvers also should be included in the structural demonstration to verify the max limit load capability or design of the rotor and helicopter system

43. Asymmetrical Flight • Combined LFs (vertical, lateral, and longitudinal) and simultaneous rotational accelerations combine to establish the max LFs during asymmetrical flight operations • Power-on, power-off, and rotor speed variations also are applicable to asymmetrical flight maneuvers • Asymmetrical flight maneuvers include rolling pullups, sideslips and yaw, and sideward flight • A rolling pullup maneuver consists of a pullup while the helicopter is in a rolling flight attitude • The roll or bank attitude results in a LF nz=1/cosΦ • The LF then is increased in proportion to the pullup and turn rates with the additional normal loads applied as shown in Fig 4-10

44. Asymmetrical Flight • The primary LFs encountered in the rolling pullup maneuver are in the normal vertical direction • However, the longitudinal and lateral loads also are significant, as is the lateral moment required for the roll rate, and all shall be considered • The rolling pullup maneuver is one of the principal design maneuvers of MIL-S-8698

45. Asymmetrical Flight • The sideslips, yaw, and sideward flight maneuvers provide the highest lateral load factors • Limit LFs for the tail boom and directional control system also are established by these maneuvers

46. Gusts • It has been established that the reaction of a helicopter to gusts is less severe than is the reaction of a fixed-wing aricraft • However, LFs on helicopters due to gusts are not insignificant • An incremental LF of 0.0 due to a gust was experienced by a commercial helicopter during a test in the Chicago area • The trend toward higher airspeeds for helicopters makes gust criteria more significant because gust LFs normally increase with airspeed • A gust load criterion is given in MIL-S-8698, including definition of a gust alleviation factor

47. Conditions Requiring Gust Load Factors • The factors that enter into the determination of gust LFs include: • RFP and applicable specifications • Fatigue considerations • Airspeed range • Configuration • Exposed surfaces • MIL-S-8698 requires the gust velocity to be considered at SL is 50 fps • FAR Part 29 specifies a gust velocity of 30 fps for design Asymmetrical Flight

48. Design Limit Landing Requirements • Consideration of both level and asymmetric landings – with and w/o forward speed – at varying helicopter sizes, weights, and configurations is essential • The parameters of the landing gear design or configuration sensitive to the anticipated operational factors are likewise of interest

49. Symmetrical Landings • Symmetrical landings can be accomplished in more than one manner – from hover; during a power-on approach with forward velocity, possibly including drift; fully autorotational with or without forward speed, and possibly with drift; or, in rear cases, under emergency conditions involving abnormal descent velocity • These rare, emergency landings usually are attributable to battle damage, vehicle malfunction, adverse weather or terrain conditions, or pilot error

50. Symmetrical Landings • Power-on landings normally result in relatively low descent velocities at ground contact, usually not exceeding 5 fps • Fully autorotational landings rarely are performed with large multiengine helicopters but are frequent with the smaller single-engine models, to the point that they shall be considered as normal landings • These normal autorotational landings usually do not exceed 6.5 fps descent velocity at ground contact • In battle zone operations, a descent velocity of 8 fps may be considered a normal sinking speed at ground contact • Under emergency conditions, or in unusual situations as noted, descent velocities of 15 fps or more occasionally may be encountered