AIRWORTHINESS: LOADS SPECIFICATIONS

Transcription

1 1 JEST-M, Vol 3, Issue 2, July-2014 AIRWORTHINESS: LOADS SPECIFICATIONS Air Commodore(retd) Professor & Head Department of Aeronautical Engineering M V J College of Engineering, Bangalore, India Abstract- Specification of requirements are intended to ensure the structural integrity throughout the operational life of the aircraft. Load Specifications, Repeated Load Specifications and Specified Aeroelastic Requirements are addressed herein. Loading cases and load distributions are covered. Introduction The flight loading cases include symmetric and asymmetric flight maneuvers and atmospheric turbulence. Symmetric flight cases are concerned with the design requirements for the strength of airframe when it is loaded in the plane of symmetry. The greatest loads arise while pulling from dive or when in a turn maneuver. A steady pitching velocity has to be considered at all points on or within the flight envelope. The pitching acceleration is to be superimposed on the normal acceleration. The tail areas of aircraft get affected the most. The available pitch authority needs to be considered for realistic description, and control criteria/control system demands can complicate the specification requirement. The conventional way to describe the symmetric flight loading is to consider a flight envelope of forward velocity and normal acceleration. The speed when taken as equivalent air speed makes such a velocity versus load factor (V-n) diagram independent of altitude effects. The compressibility effects become invisible and therefore Mach number and altitude based envelope for normal acceleration i.e. `g values become important. The ultimate load factor i.e. max n`g in certain extremely rare circumstances may be exceed by assuming that the probability of doing so is less than extremely remote say 1 in The loading cases in the asymmetric plane are concerned with rolling, yawing, and side slip motion in combination with symmetric loads. Roll pull-out for various n`g values, yawing/side slip, steady heading side slip, engine failure cases and asymmetric gust are considered under various loading conditions. The aircraft response in an asymmetric maneuver can get complicated by the coupling which can occur between all six degrees of freedom. In case of roll around velocity axis, resulting loads can be considered in isolation from the yawing motion. Loads arising out of rolling, yawing / side slip maneuvers are explained herein. Fins and rudder loads, and corresponding lateral and yaw accelerations, side slip to reach maximum over-swing are highlighted. Engine failure cases, atmospheric turbulence and gusts are explained. Differential tail plane loads form part of asymmetric flight loads[1,8].

2 2 JEST-M, Vol 3, Issue 2, July-2014 Ground loads are those occurring when aircraft is in contact, or makes contact with the ground. Aircraft configuration, high lift devices and thrust reversal have bearing on ground loads. Although the stipulations of various design loads are expressed in terms of the alighting gear units or their components there is an implicit effect upon the airframe as a whole. The energy absorption characteristics are associated with the vertical descent velocity of the aircraft at landing impact. The aircraft movement on ground forms part of ground loads. Ramp mass, take-off mass, landing mass and emergency landing mass are the four conditions for design. Three point and two point landing are further considered for distribution of energy between nose and main landing gear units depending upon the attitude of aircraft at touch down. Conventional landing is a two point landing. Regardless of how many main gear units are actually employed, an attitude where only the main gear units initially touch down is always referred to as a two point landing. A three point landing is a hard landing since attitude of aircraft is nearly level and thus landing speeds are high. The satisfactory behavior of an aircraft under single occurrence of the maximum design load may be ascertained with an acceptable degree of accuracy, however the design life of the airframe should be established for fatigue loading. The repeated load data is represented as load spectra either in diagram or tabular form. The spectra are in terms of a design condition, such as maneuver acceleration or gust velocity, which is reached or exceeded in a given period, specified in terms of number of flights, time or distance traveled. Load spectrum can be regulated through fatigue meters kept near center of gravity of aircraft. Landing gear loads, unevenness of ground surface, overload operating conditions, hard landing conditions, buffeting and noise are considered towards repeated loads. Significant turbulence occurs when one aircraft flies in the wake of another, especially when flying for considerable time during flight refueling. All these aspects needs to be considered at the design stage. The significance of the structural distortion of the airframe under the applied loading is dealt through aeroelasticity. Favorable wash-in effects are possible through interior ply arrangements inside wing for load alleviation. Interactions between aerodynamic loads and stiffness, or elasticity give rise to aeroelastic requirements. Aeroelasticity covers all interactions of aerodynamics, structure and inertia including the impact of these interactions on control and stability. The structural deformations may be either static or dynamic and hence it is necessary to consider damping effects as well as the stiffness contributions from the aerodynamic and structure sources. Specified aeroelastic requirements are addressed in this text. The aeroelastic requirements specified in the various airworthiness documents are stated in terms of speeds below which catastrophic events must not occur. These events include flutter, loss of control, and aero-servo-elastic instabilities. Torsional stiffness of wing is the most used as a design requirement to ensure adequate wing aeroelastic performance. Technical Specifications The loads experienced by aircraft fall into two broad categories, Figure-1 to refer. Whether the loads are due to maneuver or due to environmental effects,

3 3 JEST-M, Vol 3, Issue 2, July-2014 for the purpose of structural design they have been dealt in two ways. Firstly there is the limit load condition. Secondly, there is the load spectrum or the set of loads of varying magnitudes experienced by the airframe throughout its life. Limit load is the actual maximum load of a particular case anticipated to occur in the prescribed operating envelope. It is the maximum load for a particular maneuver or environmental condition and represents the most severe isolated intensity of a load case. Conventional manned aerospace vehicle structures are designed using the concept of factors superimposed on the limit load. First, there is a proof factor which is numerical value of for military and 1.0 for civil aircraft. Secondly, there is ultimate factor of 1.5 for all types of aircraft. Occurrences such as flying through severe turbulence, heavy landing, cases of emergency landing warrant inspection for structural checks for ascertaining extent of deformity; and the design must cater inspection windows and possibility of structural repairs/reinforcements. Ultimate factor is effectively a safety factor on limit load. The structure must be capable of resisting the ultimate load, and the civil requirements especially state that it must be possible to withstand this load for three seconds without collapse. For some specific cases, higher ultimate factors are stated. The layout of the design is kept in the interest of keeping such factors least applied. The application of proof and ultimate factors cover the limit load condition of a particular case. Other measures are necessary to safeguard the integrity of the structure when it is subjected to repetitions of loads. The structure should be designed to have an estimated life, or safe life of three or more times that is actually intended in service. This is demonstrated through Full Scale Fatigue Tests (FSFT) for the predicted loads. Predicted loads need verification through strain gauging of the prototype. Military combat aircraft are designed using the safe life approach. The number of landings, flying hours and the calendar life form the criteria for structural integrity. The fail safe concept is critical in flight control system where electronic flight control is essentially used for statically unstable airframes. Complex sets of requirements and objectives include specifications of aircraft performance, safety, reliability and maintainability, subsystem properties and performance. Alternatively, a probabilistic design approach is seen to ensure the continued integrity of the structure. The concept of factors superimposed on a limit load are replaced by a statistical demonstration of the required failure probability. It is particularly applicable in some circumstances, namely - (1) when the experience lacks to especially realistic factors, (2) when the randomness in loading is high or material properties are a variable, or (3) case of complex systems like electronic flight control system. Damage tolerance of structure is intended for fail-safe operation i.e. to ensure should serious fatigue, corrosion, or accidental damage occur within the operational life of the aircraft, the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the damage is detected. The fail-safe evaluation should encompass establishing the components which are to be designed as damage-tolerant defining the loading conditions and extent of damage, conducting sufficient representative tests and/or analysis to substantiate the design objectives i.e. life-to-crack initiation, cracks propagation rate and residual strength. Design features for attaining damage tolerant structure includes : (1) multiple load-path construction and use of crack

4 4 JEST-M, Vol 3, Issue 2, July-2014 stopper to control the rate of crack growth, and to provide adequate residual strength, (2) after initiation of cracks, provide a controller slow rate of crack propagation combined with high residual strength, (3) detection of failure in any critical structure element before the reserve factor (RF) reduces to a value of unity, and (4) provision to limit the concurrent multiple damage. Safe-life strength evaluation method intends to ensure that catastrophic fatigue failure as a result of repeated loads does not occur. Under these methods, loading spectra should be established. The fatigue life of the structure for the spectra should be determined and a scatter factor should be applied to the fatigue life to establish the safe-life for the structure. The loading theoretically estimated or established through wind tunnel tests need to be correlated with flight load through strain survey. In the interpretation of fatigue life analysis, the effect of variability should be accounted for by an appropriate scatter factor. Recorded load and stress data entails instrument the aircraft in service to obtain a representative sampling of actual loads and stresses experienced. Current Airworthiness Codes. The civil aircraft requirements are seen to have considerable commonality between the European Joint Airworthiness Authority (JAA) and equivalent FAA of US. The specification requirements must be met at each appropriate combination of weight and center of gravity within the range of loading conditions for which certification is requested by tests upon the aircraft or by calculations based on, and equal in accuracy to the result of testing. Load distribution, ranges of weights and center of gravity within which the aircraft may be safely operated must be established. If a weight and center of gravity combination is allowable only within certain load distribution limits (such as spanwise) that could be inadvertently exceeded, these limits and the corresponding weight and center of gravity combinations must be established. Accelerations are experienced by aircraft when it encounters variations in conditions in the air such as changes in wind direction, velocity and turbulence. Sudden changes are known as discrete gusts and velocity variations of over 20 m/s equivalent air speed (V EAS ) have been measured. The design speed for the maximum gust intensity affects the `-g boundary of flight envelope the most. It may be chosen to provide an optimum margin between the low and the high speed buffet boundaries. JAR provides the minimum gust speed based upon the incremental load factor resulting from the aircraft encountering a gust. The definition of gust speed in Def. Stan is somewhat similar but does depend on whether the maximum Mach number in horizontal flight is greater or lesser than unity. For the Mach number less than unity and if there is no weapon system, the gust speed shall be either the speed determined by the intersection of the line representing the maximum lift coefficient and the 20m/sec gust line on V-n diagram or related with the incremental load factor resulting from gust. In case of weapon system aircraft where maximum Mach number is one or higher, the gust speed is determined by mission requirements. The horizontal and vertical tail surfaces and their supporting structures must be designed to the load conditions resulting from the aircraft encountering the specified gust velocities in direction at right angles to the flight path. Due account must be taken of the stability and control aspects considering failure modes like trim runaway etc.

5 5 JEST-M, Vol 3, Issue 2, July-2014 Flight loading cases. Steady pitching and pitching acceleration cover the n`g loads. The handling qualities of aircraft must be demonstrated at all the corners of V-n diagram. The specification of pitching acceleration over and above the normal acceleration depends upon control, inertia and aerodynamic characteristics of the aircraft. Figure 2 shows the combined pitch and normal acceleration plots for a small transport configuration. The loading in asymmetric plane is due to rolling, yawing and side slip motions. The combined loading as a result of pitch and roll needs to be established in the case of aircraft with highly swept wings configurations. The roll performance requirements are based on time to change in bank angle, roll rate and rollmode time constant. In case of critical engine in failed condition, civil requirements specify that at low speeds with flaps in takeoff position and undercarriage retracted, it must be possible to roll the aircraft to a steady banked turn of specified value, both with and against the failed engine. The differential thrust takes away the rudder travel availability. The loads on rudder to hold straight and level flight condition increase. Roll performance requirements at low speeds are to rapidly lift the wing during landing approach in case of wing drop. Roll pull out and wind-up turns involve combined roll and pitch maneuver. Specific loading conditions are prescribed when a rolling maneuver is combined with a pitch maneuver. The roll and pitch effects are analyzed separately and symmetric pitch part superimposed. The additional effect of roll motivator is then added. Roll motivator deflection requirements are speed dependent e.g. at speed V D, the deflection requirement is 1/3 of steady roll rate which occurs in condition V A. The yaw control is considered fixed at trim or deflected to avoid side slip. The effect of flap to ease out maneuvers and use of air brakes etc. are considered for analysis. Roll motivator deflections are specific e.g. a value of n 1 g is taken for transport aircraft. Yaw motivator requirement are determined by the performance requirements i.e. steady heading in case of single engine on a twin engine aircraft or critical engine failure in four engine aircraft in the presence of cross wind. Yaw motivator requirement are determined by the performance requirement for a case of steady heading with single engine for a twin engine aircraft, and for a critical engine failure in a four engine aircraft with the presence of cross wind. In case of two engines failed on same side, the aircraft should be able to maintain heading at lower specified altitude with reduced power settings. The rudder and elevon are interconnected through a coupling in case of fighter / attack airplane so as to limit the over movement. There are other considerations which may assist in determining rudder deflection limitation e.g., a) it must not be possible to stall the fin dynamically as a consequence of rudder application at all angles of attack. This is crucial for entry into spin. As a thumb rule, the fin dynamic stall angle shall be up to 1.5 times the static value. The strake vortices created from sharp leading edge devices help in delaying the rudder stall at high angles of attack. Appropriate combinations of normal accelerations with lateral and yaw rates and accelerations are specified (Figue- 3 to refer). Instantaneous rudder deflection for `over swing angle case must be examined for the resulting dynamic motion). The value of max equal to 1.5 times the resulting yaw ( ) because of steady heading side slip (SHSS) is normally considered.

6 6 JEST-M, Vol 3, Issue 2, July-2014 This corresponds to maximum lateral velocity. The doublet input to rudder resulting in yaw oscillations shall be examined. When the rudder is moved in oscillatory fashion at the damped natural frequency of the aircraft in yaw, the resulting motion is a `fish-tail maneuver. The resulting vertical fin loads will continue to increase with time and it is necessary to prescribe limit to motion. The limit is normally taken as one-and-a-half cycle for fighter / attack aircraft. Yaw motivator induced lateral maneuver cases must be analyzed. The deflections of the yaw motivators are determined by several considerations and the most-critical case is worked out. Initial flight is considered to be straight and level. The loading is to be analyzed for all altitudes and for minimum speed to design speed. The step and doublet input to yaw control are considered. The maximum permissible step input either as a control limit or pedal force limit is injected. For fighter / attack aircraft, yaw control is moved sinusoidally at the natural damped frequency of the aircraft in yaw. The maximum rudder deflection is to be twothirds of maximum permitted throw and the input is to consist of one-and-a-half cycles of the pilot s inceptor (i.e. left-central-rightcentral-left-central). The rolling motion caused by the yaw motivator must be arrested if it is beyond a prescribed limit / maximum dynamic side slip value. Any pitching as a result of yaw control input, must be stopped if it exceeded by ¼ `g increment. The sinusoidal rudder input should be done at low angles-of-attack since the resulting dutch motion can become divergent spiral and uncontrollable at medium and high angles-of-attack. Horizontal stabilizer loading in yawed flight needs to be analyzed. UK military requirements suggest to consider C mo (i.e. `o lift pitching moment coefficient) having increased by / w.r.t. straight and level flight. In case of multi-engined aircraft, the engine failure could result in additional loads on control to maintain heading since the loss of moment arm could be significant. It is essential to ensure that the loads resulting from the failure of engine powers do not exceed design loads of airframe / control column limitation / adequate rudder power availability to correct swing and maintain straight path. In case of one engine failure, it should be possible to maintain straight flight through rudder trim setting. Trimrunaway should be taken into account. Gusts and atmospheric turbulence have cumulative effect in affecting structural life. Discrete gusts of 20m/sec are considered towards influencing the lift boundary. A m/sec gust is considered towards V H definition. The `-g boundary of V-n diagram while flying with max reheat power gets influenced by this consideration. A slow down on transport aircraft alone may result in a design limitation in the presence of gust. In case of class IV aircraft, diving speed boundary gets affected by specified gust value of 7.60m/sec only. Gust asymmetry shall be considered for transport aircraft. Lateral gusts aligned between the vertical and horizontal directions must be considered. Continuous turbulence is modeled through Power Spectral Density (PSD). Load Specifications For each of the conditions, it is necessary to interpret the relevant flight maneuver load cases. This can be best dealt by considering firstly, the aircraft in trimmed level of flight resulting in steady flight loads and secondly, the increment in loading as a consequent upon moving the control input. The load specifications towards maneuver, atmospheric disturbance

7 7 JEST-M, Vol 3, Issue 2, July-2014 and the ground operation are explained herein [1-2]. Maneuver Loads. There are three basic ways of moving a control motivator, as below : a) Unchecked Control is moved and held at the new position. Such a movement is required to overcome stability of vehicle. b) Checked Where the control is deflected and brought to neutral position. Such a movement is required for a case of neutrally stable airframe. c) Excitation Where the control is moved with a oscillatory input. Such a movement input is required for case of tracking or in case of formation flying. The specification of pitching acceleration is worked out for a maximum deflection of pitch control while the aircraft is in trimmed straight and steady level flight at V A. For example, the prescribed minimum design pitching acceleration for a transport aircraft case as per JAR is as below : (a) Nose up.. = 39 n max (n max -1.5)/V The loads arising from maximum attainable elevator movement in a unchecked maneuver must be analyzed. In the case of step input to the control, one of the design maximum occurs at the instant the control is applied. This is because the rate of application of control is significantly greater than the response of the aircraft. The initial maximum load will be experienced early in the subsequent motion. Although the unchecked mode of the pitch control application has only to be applied at the speed V A, a comparison of loads arising with those in a checked maneuver shall be made for certification purposes. Figure 4 shows nose-up pitch from level flight due to elevator up step at V C of a transport aircraft. Sinusoidal application of pitch control be analyzed for varied phase differences from un-damped natural frequency of short-period modes. Current flight system of class IV aircraft provide 6 db gain and 135 o phase margin for the input-output response for which the loads must be ascertained. Loads on the horizontal stabilizer should be examined for : a) The initiation of pitch maneuver, and b) load arising due to local angle of attack on stabilizer surface. The chord-wise loading of these two cases is different and consequently the torque loading is very different. It may also be not necessary that maxima of each effect are coincident in terms of time. Loads on leading edge and trailing edge surfaces be analyzed for three conditions : a) angle of attack, b) control surface deflection, c) asymmetric effects such as rolling. The yaw motivator loads are considered from two inputs, (a) the step input i.e. unchecked, and (b) the oscillatory input. Loads arising out of maximum in a steady heading side slip and the design maximum total load on the vertical surface is usually given by the over-swing condition. Rudder input in a coning motion can be a more severe case when used during rolling motion. Pure rolling motion i.e. roll around body axis generates maximum tensile loads on wing mounts. The load factor excursion is maximum. This flick roll motion is usually a prohibitive one. It generates higher vertical fin loads than those produced on fin in SHSS. The magnitude of lateral and yaw accelerations is needed in order to evaluate the inertial relief effects in the lateral motion. It is also desirable to consider the lateral accelerations experienced by the crew to ensure that these are within tolerable limits. The air intake performance under the flick roll and angle of

8 8 JEST-M, Vol 3, Issue 2, July-2014 attack ( ) combinations as well as roll and - combination should be ascertained. It is possible that there may be need to limit the lateral acceleration for reasons of crew tolerance or air intake gearing affected and to provide a corresponding limit to the allowable deflection of the yaw motivator at higher speeds and hence a limit upon load. The asymmetric horizontal stabilizer load due to side slip is quite complex since it depends upon the geometric and aerodynamic details. MIL specifications entails to consider that there is a lift coefficient difference on the two halves of the tail plane of 1.0 or C Lmax whichever is lower. Differential tail plane for roll compliance in fighter class of aircraft have design specific differences on two halves of tail plaine. Some load cases are graphically shown in Figures 5-7 for a typical small aircraft. Loads due to Atmospheric Disturbances. The structural loading arising as a result of aircraft encountering gust and turbulence is explained herein. The dynamic response of the airframe in a discrete gust encounter is an important consideration as it can significantly increase the structural stresses. While the gust is considered as a column of air moving at some speed, the continuous turbulence is specified in terms of PSD. Alleviated sharpedged gust concept using a gradient distance of 12.5 of mean aerodynamic chord (m.a.c.) of wing formed the basis of discrete gust load requirements for a considerable period and is still being used as design criteria. The alleviated factor is applied over the discrete gust since the discrete gust so called `sharpedged gust is unrealistic as it does not represent the boundary effect and its application to the aircraft results in a higher acceleration increment than would occur in a practical situation. It was also realized that the assumption of a single, typical, gradient distance is an over-simplification and this led to the `tuned gust concept, that is, one where the gradient distance used is that which gives highest acceleration increment for a given design gust velocity. In practice it has been found that the effect of variation in the gradient-distance on the acceleration increments is not large, but can have an important impact upon the fatigue life assessment. The corrected design velocity of the elevated sharp-edge gust is FU de, where F is the alleviating factor and U de is the design gust velocity considering (1- Cosine) vertical gust, the heave effect is worked out with allowance for following two effects. a) The lag in the build-up of lift consequent upon an instantaneous change of the angle of attack. b) The lag due to the fact that the angle of attack across the chord does not change simultaneously, but gradually as it progresses into the gust. For a given aircraft both these effects are a function only of the distance the wing travels into the gust. While the alleviated sharp-edge gust may only affect the heave motion of the aircraft, the aircraft with canard / foreplane controls shall experience pitch up. The gust V-n diagram is worked out by plotting the function (n G +1) against the forward equivalent air speed. In case of wings at low fuel weight, the gust load effects are more severe since wing loading is lower. It must be noted that the wing-body load is (n G +1) mg so that a high value of n G often associated with a low wing loading case may not actually give a critical load. When the vehicle is subjected to compressibility effect the n G is not directly linear with forward air speed. At supersonic speeds the reduction of lift curve slope with increase in Mach number may cause a lower n G, even though of the apparent direct

9 9 JEST-M, Vol 3, Issue 2, July-2014 increase due to higher Mach number which relates to high true and not V EAS. Ground loads. The vertical sink velocities effect the ground loads the most. Aircraft movement on ground at maximum `Ramp mass and the energy absorption as a result of landing impact influence the energy absorption characteristics of the undercarriage. MIL-A-8862 supplemented by MIL-L-8713 covers wide range of design conditions and all possible configurations. The three values of mass i.e. `Ramp mass, `Take-off mass and the `Landing mass influence the aircraft design mass conditions. In case of flare at touch down, only the main gears absorb the energy at touch down. In case of `three point landing, the nose gear shares the energy absorption. Since the actual energy absorption requirement is vertical absorption parallel to vertical plane of c.g., it is convenient to describe the characteristic of aircraft landing shock strut in terms of the vertical load and vertical axle travel. The energy to be absorbed on impact is the sum of kinetic energy due to the vertical velocity at the instant of impact and the potential energy. The potential energy is equal to the product of the weight and the vertical displacement of the unit occurring over the period of time from the instant of impact to that when the shock absorber and tyre have reached their maximum deflections. Def. Stan specify that at least 67% of the energy must be dissipated in the initial closure and rest dealt with on the first rebound. Vertical velocity during landing should cover the limit as well as the ultimate energy condition which is the takeoff mass. Neither the shock absorber nor the tyre must reach maximum deflection under the limit load or maximum reaction. One of the two may reach maximum deflection in the ultimate energy case. In many design codes the application of an ultimate vertical velocity is taken with a factor of 1.2 at the landing mass condition. Since it is an ultimate case the load is factored by unit, the usual proof and ultimate factors do not apply. MIL-A-8862 specifies alternatively an overload landing case which is that at mass of 1.15 times the design landing condition associated with a vertical velocity of 93% of corresponding design value. The greater of these vertical loads are used as datum values in specifying the overall loading on main under carriage units in twopoint landing case. The main gears loads are inevitably located aft of center of gravity of aircraft, the consequence is a nose-down pitching moment. This results in the eventual impact of the nose-wheel so that the unit absorbs some of the rotational energy and a vertical load is developed. In a three-point landing condition where the nose-wheel contacts the ground at the same time as the main wheels and the nose gear unit absorbs a part of the vertical energy. As the aircraft touches the ground the main wheels must rapidly `spin-up to the rotational speed equivalent to rolling at the forward speed of aircraft. This results in a horizontal friction force which coupled with vertical reaction creates strong nose pitch down moment. Vertical loads are associated with drag loads and side loads as well as. These are as below: 1 a) Landing with drag alone Case 1(a). b) Landing with drag and side load - Civil requirements Case 1(b) c) Landing with drag and side load - Military requirements Case 1(c) 2. Side load case Load Case (2) 3. High drag Load Case (3) 4. One side landing Load Case (4) 5. Rebound of un-sprung Load Case (5)

10 10 JEST-M, Vol 3, Issue 2, July-2014 Load case (1a) is the basic case specification of the combined drag load, and the side load varies with different sets of requirements. The combination of full resulting vertical action and side load is not always required. While Def. Stan prescribes full value of vertical reaction, other requirements are seen to allow a lower value of 75%. Load case (3) is for a spin-up loading condition and spring back. The fore and aft force is dependent upon the time it takes for the wheel to `spin-up, and the `spring-back effects from the strut. Bending strain energy in the leg causes a forward springing of the oleo, known as `spring-back. The springback loads are calculated for ground friction coefficients of up to an average value of 0.55 at touch down speeds of at least 1.2 times the stalling speeds. One-side landing may occur due to several reasons. The loads in this case (No.4) are not normally a design case for strut design but may be a critical load case for airframe structure between the main under carriage gear units. Main landing gears can be arranged in such a way that loads in an asymmetric landing are shelved and are not higher then the symmetric case. Load case (No.5) is for the rebound of un-sprung parts. Attachment of the un-sprung mass (i.e. wheel, axle, and lower part of the shock absorber mechanism) should be designed to withstand a limit load factor of 20 along the axle of oleo. Load cases resulting from ground maneuver conditions (braking, reverse braking, turning, pivoting, towing), operation of uneven surfaces / unpaved, partially graded surfaces must be taken into account. Phenomenon known as `wheel barrowing may occur if nose wheel is touched firmly before other wheels touch the runaway. Wheel barrowing may be described as an attitude or condition in tricycle gear equipped aircraft that is encountered after initial ground contact during landing rollout, wherein the main wheels are lightly loaded or clear of the runway. However, the nose wheel is firmly in contact with the runway thus causing the nose gear support a greater than normal percentage of aircraft weight while providing the only means of steering. In a crosswind, the airplane in this situation tends to pivot rapidly about the nose wheel, in a maneuver very similar to a ground loop in a tail wheel type aeroplane. Other indications of `wheel barrowing are wheel skipping and/or extreme loss of braking effect when the brakes are applied. Normally, `wheel barrowing may be encountered if the pilot is utilizing excess approach speed in a full flap configuration that results in the aircraft touching down with little or no rotation. After this touch down, the pilot may then try to hold the aircraft on the ground with forward pressure on the control wheel. Under these conditions, braking and steering capability is severely diminished and `wheel barrowing is likely to result. `Wheel barrowing accidents have occurred during crosswind landings made by pilots flying aircraft equipped with stabilizer type elevators and nose wheel/rudder steering, and utilizing the `slip technique for crosswind correction. On most general aviation aircraft, the nose wheel steers when rudder is applied and, for this reason, such landings require careful rudder operation just prior to and during touch down. The `slip method of drift correction is favored by the majority of pilots as it accomplishes the desired results without presenting the need for a last minute directional correction prior to touch down. A corrective action must be based on a number of factors, i.e. degree of development of the wheel barrowing, pilot proficiency, remaining runway length and aircraft performance versus aircraft configuration. Only after considering at

11 11 JEST-M, Vol 3, Issue 2, July-2014 least these factors, the pilot should initiate corrective measures. Specification of Repeated Loading Individual loadings on airframe cause `fatigue damage over the period of time and collectively determine the life of airframe. Load spectrum should be based on measured statistical data of the type derived from load history studies and where data is insufficient, conservative estimate of the anticipated use of aircraft be made. The structural damage caused by a given increment in load is not only a function of its magnitude but depends also upon such factors as the initial condition to which the load increment is added. Loading on individual airframe components, air-load distributions, specification and analysis of repeated loading are explained below. Loading on Airframe Components. During its specified life any aircraft structure or system is subjected to a history of load fluctuations which occur during the various ground and flight operations. The loads comprise of flight loads (Maneuver and gust), ground loads (taxing, landing, impact, turning, engine run-up, breaking and towing), and pressurization loads. Loading history for each phase of operations is then reduced in to individual cycles. These loads are either concentrated or distributed depending upon the structural arrangements e.g. the longitudinal acceleration / breaking do not cause a significant overall loading of the airframe but affect the local components i.e. the attachment of point loads. Most severe case of deceleration condition occurs with negative thrust through thrust reversals with emergency breaking. These are limited to not more than 0.65 `g. More severe loading cases occur when the aircraft is required to operate under assisted take-off and arrested landing conditions. Typical arrester gear deceleration are of the order of 4`g to 6 `g. Considering proof and ultimate factors, the fatigue load test value could be around 10 `g. Loads arising as a result of abrupt control inputs needs to be ascertained. For example, in the case of civil transport aircraft abrupt pitching maneuvers when applied to structural load analysis are maneuvers involving a single rapid application of the elevator in a prescribed manner. Generally, two types of abrupt pitching maneuvers need to be considered: a) abrupt unchecked elevator maneuver at V A speed, and 2) elevator checked maneuvers at V A and V D speeds. In the first case the elevator is suddenly moved to obtain extreme positive pitching acceleration (nose-up). Transient rigid-body response of the aircraft must be taken into account in determining the tail load. Aircraft loads that occur after normal acceleration at center of gravity exceeding the maximum positive limit maneuvering load factor need not be considered. In case of checked maneuver, rational pitching control motion versus time profile must be established in which the design limit load factor will not be exceeding. As per JAR (c) (2), checked pitch maneuver must be analyzed for nose-up conditions to the maximum design limit load factor and for nose-down conditions to a load factor of zero. In case of checked maneuver, while flying in steady flight condition at any speed between V A & V D, the pitch control is moved rapidly in sinusoidal motion before return to trim. Military specifications require a trapezoidal shaped elevator input that is checked back to neutral for conditions with free stream center of gravity and beyond neutral (or the trim point) to 50% of the original input elevator. The rolling maneuvers are considered with specified entry normal

12 12 JEST-M, Vol 3, Issue 2, July-2014 accelerations to cover the roll pull outs in case of class IV aircraft. In case of yaw, two types of yaw maneuvers are considered for design : i) Max rudder in level flight SHSS, and (ii) engine out condition, whereby abrupt application of the rudder is made in conjunction with the resulting side slip due to unsymmetrical engine thrust. The military criteria for rudder maneuver requirements for transport aircraft are different than that of civil aircraft, in that, that basic maneuver conditions are the same, but the definition of the available rudder is different for the over yaw and steady side slip condition. The loading on lifting surface is distributed type. The design condition for wing derives from the superimposed upon the loading in the trimmed conditions. The loads due to symmetric and rolling maneuver as well as from high lift devices and lift spoilers / dive brakes and control surface movement form part of the load distribution. Specific loading conditions are prescribed when rolling maneuver is combined with a pitch maneuver e.g. roll pull-out entry normal acceleration. The two effects are generally analyzed separately and the results superimposed in appropriate proportions. Roll results in side-load on vertical fin. Side load is applied independently of other conditions and must not be less than a load factor of 1.33 or 0.33 n 1, whichever is a higher value. Vertical stabilizer gets maximum loading in a steady heading side slip motion. Air turbulence forms distributed loading on lifting platforms and vertical fins. In case of some military transport aircraft, the pressurized compartment may be limited to the region of crew occupancy only. In case of cabin pressurization, two standards of cabin pressurization are laid down i.e. low differential and high differential pressure. In case of low differential pressure, maximum cabin altitude of 6.7 km is allowed and in the case of high differential pressure, maximum cabin altitude allowed is 2.5 km. While testing the cabin for fatigue the working differential pressure should be at least 1.5 times the specification value and aircraft speed is to be the design speed. The loads coming on power plant mount are following : i) Thrust (forward and reverse) ii) Engine torque (including seizure case) iii) Gyroscopic couples due to the angular motion of aircraft iv) Inertial forces (linear & angular accelerations) v) Air loads on nacelles, slip stream effects vi) Thermal effects Small aircraft gyroscopic couples load requirement must be considered with symmetric and asymmetric maneuvers and gust cases. Design specifications include aircraft rotation rates and accelerations in addition to the usual maneuver and gust cases. These include all possible following combinations ( e.g. case of non-aerobatic category as per JAR ): i) Yaw rate of 2.5 rad/sec ii) Pitch rate of 1.0 rad/sec iii) Normal load factor 2.5 iv) Max continuous thrust Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner. Dutch excitation with rudder doublet application is typically shown in Figure-7 for a case of class-ii aircraft. The civil requirements specify accelerations to be assumed parallel to the hinge-lines of auxiliary surfaces for the purpose of designing the hinge brackets. For vertical surface the acceleration is 24 `g and

13 13 JEST-M, Vol 3, Issue 2, July-2014 for horizontal surface the acceleration is 12 `g (JAR to refer). Air-load Distribution. Overall air load on the various airframe components of the aircraft result in stressing information. The torque on a chord wise section is the moment of the chord wise air load about a reference point. Care must be taken to distinguish the total pitching moment due to chord wise loading consequent upon the deflection of the flap and flap hinge moment. Loads need to be distributed across the relevant components. Vortex shedding from the fuselage at high angles of attack influence the loading on lifting surfaces and vertical fin. The high angles of attack conditions essentially occur at subsonic flows. Experiments on inclined bodies of revolution at supersonic speeds have shown that a pair of vortices do form for all but very small incidences. These vortices separate from the upper surface and follow on approximately free stream direction in a similar manner to the trailing vortices of a wing. The drag associated with this cross flow velocity is effectively equivalent to the normal force on the body 2 i.e. (V ). The effect of fuselage on a wing and that of wing on body results from interference velocity potential ( i ). Elliptic horizontal body shapes have favorable interference effects as compared to any other form. The effect of body on wing is due to two factors. Firstly, at the wing body junction the body usually has a relatively thick boundary layer and this has the effect of reducing the pressure changes on the wing so that the resulting loading at the root of the wing gets reduced. The second effect is that the body usually induces up wash field over the wing causing an increase in normal force. The region of the interference of the wing on body is governed by the transonic area rule. At the supersonic speeds, it is marked by tracing the helices made by the Mach lines from the wing root chord on the body. In ideal flow, the chord wise pressure distribution at the root of the wing is transmitted to the body in the region aft of the position of the helices from the root leading edge and decreases in magnitude with the distance aft from the root chord. The effect of the body boundary layer is to smooth this effect, especially at some distance from the root, but the distribution of the load is essentially similar for cases where there is no marked separation over the body in the region of the wing. The effects of repeated loading and environmental exposure on stiffness, mass and damping properties should be covered for verification of integrity against flutter and other aeroelastic mechanism. Combined load cases from each category are worked out though in any given category each load being lower in magnitude than the corresponding limit deign case. These loads individually cause `fatigue damage and hence collectively determine the airframe life. The structural damage caused by a given increment in load is not only a function of its magnitude but depends also upon such factors as the previous loading history and the initial conditions to which the load increment is added. The specification requirements are specified to ensure the integrity of airframes during service life and demonstrated under fatigue loading by analysis and satisfactory demonstration and in-flight measurements of loads. The magnitude and frequencies of combined loading and number of occurrences as well as that of individual loading occurrences is taken into account. There is some commonality in the design process for the safe life and fail-safe concepts i.e. the crack initiation as well as

14 14 JEST-M, Vol 3, Issue 2, July-2014 the crack growth are key factors in both the cases. The fail-safe design has an alternative load path. In case of safe life concept the life is inclusive of an appropriate load factor. Fatigue (safe-life) evaluation involves in estimating or measuring the expected loading spectra for the structure, fatigue testing of structure, providing data for inspection and maintenance instruction. In addition, fatigue initiation from sources such as corrosion, stress corrosion, disbonding, accidental damage and manufacturing defects should be covered based on experience / expertise. In the interpretation of fatigue analysis and test data, the effect of variability should be taken into account by an appropriate scatter factor. Loads for fatigue testing should be verified through in-flight measurements. Recorded load and stress entails instrumentation of aircraft to obtain representative sampling of actual loads and stress experienced. Repeated load tests of replaced parts can be utilized to re-evaluate the established safelife. This data is also useful for life extension program through rework. In order to establish the total technical life of airframe, it is essential to consider all possible loading combinations, frequencies of occurrences and magnitudes. The subject is covered in JAR , Def. Stan Chapter 201 and MIL-A- 8866A. MIL-A-8867 prescribes the ground test requirements for assessment of airframe life. The combined loading cases in a given category can be represented on a stress (S) and N (repetition of a stress). Such a S-N curve for a given airframe component must be established against a load spectrum. The load data is to be presented as load spectra either in diagram or tabular form. The spectra are in term of a design condition such as a maneuver acceleration value or gust value i.e. reached or exceeded in a given period, specified in number of flights, time or distance traveled. Figure 8 typically shows `g versus occurrences/ hr of flight. The frequency of a given maneuver level may be obtained by tabulating the frequency of each exceedance and successively subtracting from the highest to the lowest. MIL-A-8866A gives this data for various class of aircraft. Fatigue loading data for asymmetric loads is sparse. The roll and yaw control movements generated loads are based on analytically estimated and verified from flight data. For the case of civil transport aircraft, atmospheric turbulence are of large importance. Use of an appropriate gust analysis along with knowledge of aircraft speed and altitude enables such information to be converted into a normal acceleration spectrum based on hours flown. The frequency and magnitude of lateral turbulence is considered slightly higher than vertical gusts for altitudes below 3 km. Buffeting, noise turbulence and engine wind milling are some more types of repeated loads. Extent of specification for fatigue testing involves in as to how much of aircraft need to be included in full-scale fatigue specimen. It requires evidence on several fatigue test defects and study on their locations and types. The general structural definition comprising of major components includes cut outs, ailerons and flap attachment or wing / body attachment region. The defects can be split into three categories, namely : a) these could have been represented by a simple fatigue test involving only the component with straight forward and fittings in a standard fatigue machine, b) these could have been represented in the laboratory by a special rig but probably loaded by a standard fatigue test pulsator ; within the rig it is expected that some local structure need to be represented such as a spar web with local

15 15 JEST-M, Vol 3, Issue 2, July-2014 skin and booms with loading applied possibility in more than one direction, c) the only way to detail testing this would be to load representative a complete section of structure (e.g. wing box spar or fuselage section). Another question arises as to how long one should continue with a full-scale fatigue test especially in the event of failure taking place and subsequent to structural rework. Figure 9 shows for the four different aircraft the percentage of total test defects versus times the design target life. In setting up a full-scale fatigue test, the time taken is important from a phasing aspect both in relation to certificate of airworthiness and to production rates. If one decides to build an aircraft and carryout no fatigue testing whatsoever it is considered that the stress levels for design would have to be reduced for airworthiness reasons. The fatigue sensitive structural weight would be increased by 20% to double the calculated life which is the margin usually demanded in the absence of testing. Increased maintenance and inspection would be necessary if such penalty is kept low. In addition, life extension programme involves additional fatigue testing at the end of given life by airworthiness authorities. Manufacturers consider design criteria by which the safety objectives are achieved thereby, scope exists for extension of life. Accumulation of operational experience and fatigue and fail-safe testing together with new tools in the form of probability analysis and subsequent capability and safety analyses have put the picture in the perspective. The purpose of fatigue investigation is to minimize the rate of structural fatigue failures in services and to establish suitable inspection and maintenance procedures. Fatigue investigation can not 100% define, a structure which will be free of failures for some predetermined time period. The design of relatively flexible structure, especially on large aircraft, requires earlier and better recognition of its dynamic response to rapidly applied loads during turbulence and maneuvers. Aeroelastic Specification Distortion of the airframe under the loading conditions should not affect the performance. The stiffness of structure should be sufficient so as to ensure that the airframe will not become structurally unstable under aerodynamic loading. Interactions between aerodynamic loads and the stiffness, or elasticity of the structure give rise to aeroelastic requirements. Aeroelasticity covers all interactions of aerodynamics, structure and inertia including the impact of these interactions on control and stability. The structural deformations may be either static or dynamic and hence it is necessary to consider damping effects as well as the stiffness contributions from the aerodynamics and structural sources. The dynamic response of a structure is important in two main areas : a) the dynamic stress factors which may occur when loads are applied rapidly, b) the more general interaction with aerodynamic forces with the possibility of the occurrence of flutter. Dynamic response is more complex then the static distortion. The aeroelastic requirements specified in the various airworthiness documents are stated in terms of speed below which catastrophic events must not occur. These events include flutter, loss of control, and aero-servoelastic-instability. The speeds are quoted in terms of design speeds (V D ) and vary from 1.15 x V D for Military aircraft and at least 1.25xV D for Civil aircraft. Some reduction of the margin over the design speed may be possible when an active control system is used for flutter suppression.