AIRWORTHINESS: LOADS SPECIFICATIONS · AIRWORTHINESS: LOADS SPECIFICATIONS Air Commodore(retd) S C Gupta Professor & Head Department of Aeronautical Engineering M V J College of Engineering,

1 JEST-M, Vol 3, Issue 2, July-2014

S C Gupta

AIRWORTHINESS: LOADS SPECIFICATIONS

Air Commodore(retd) S C Gupta

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

107.

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].


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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,


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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 1.125 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, sub-

system 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


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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 (VEAS) 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 25.335

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.00-970

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.


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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 roll-

mode time constant. In case of critical

engine in failed condition, civil requirements

specify that at low speeds with flaps in take-

off 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 VD, the deflection requirement

is 1/3 of steady roll rate which occurs in

condition VA. 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 0-0.67 n1g 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

òver 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.


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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 two-

thirds 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-right-

central-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 Cmo

(i.e. ò’ lift pitching moment

coefficient) having increased by -.0015/

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. Trim-

runaway 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 15.20 m/sec gust is considered

towards VH 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


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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 VA. For example, the prescribed

minimum design pitching acceleration for a

transport aircraft case as per JAR-25.331 is

as below :

(a) Nose up

..

= 39 nmax (nmax-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 VA, 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 VC 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 135o 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


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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 CLmax 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 sharp-

edged 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 `sharp-

edged’ 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 FUde,

where F is the alleviating factor and Ude 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 (nG+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 (nG+1) mg so that a high value of nG

often associated with a low – wing loading

case may not actually give a critical load.

When the vehicle is subjected to

compressibility effect the nG 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

nG, even though of the apparent direct


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increase due to higher Mach number which

relates to high true and not VEAS.

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 under-

carriage. 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. 00-970 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 take-

off 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 two-

point 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)


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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.00-970 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 spring-

back 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


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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 VA speed, and 2) elevator

checked maneuvers at VA and VD 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

25.331 (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 VA & VD, 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


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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

n1, 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 23.371):

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


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for horizontal surface the acceleration is 12

`g’ (JAR 25.393 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

i.e. (V2

).

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


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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, dis-

bonding, 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 safe-

life. 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 25.571, Def.

Stan. 00-970 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


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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-servo-

elastic-instability. The speeds are quoted in

terms of design speeds (VD) and vary from

1.15 x VD for Military aircraft and at least

1.25xVD 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.


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Structural distortion can result in a

reduction of the effectiveness of controls to

the extent that in some cases the forces and

moments consequent upon the deflection of

a control motivator are opposite in direction

to that signaled. This can happen due to

extreme aileron requirement where

aerodynamic and structural effects exactly

cancel. This may be overcome by the use of

all-moving differential tail plane where pitch

and roll effects are together obtainable from

tail plane. The spoiler action through wing

mounted surfaces assist roll at lower speeds

in such designs. Another form of static

aeroelasticity is wing torsional divergence

also known as Mach divergence. This can

occur when the twisting of the lifting surface

results in a change in angle of attack of the

local aerofoil to the extent that the further

induced twisting moment progresses to

structural failure. The speed at which

divergence occurs is proportional to the

square root of the torsional stiffness of the

wing. The most significant parameter is the

location of the local center of pressure

relative to the axis of twist of the structure

(the aerodynamic torsional moment being a

function of dynamic pressure). Torsional

stiffness is the main parameter.

Flutter is oscillatory motion

occurring either on main lifting surface or

on control surface. Main lifting surface

flutter arises primarily from a combination

of dynamic flexural and torsional distortions

of the surface but can also involve body

freedoms and other interactions such as

servo controls. Critical speed of flutter is

primarily dependent upon the frequency of

the fundamental torsional vibration mode of

the surface. Torsional stiffness plays a large

part in determining this but it is also

dependent upon inertia effects, especially

those of large concentrated masses.

Flexure-torsion flutter together can occur

especially on a high tail surface. Control

surface flutter arises as a consequence of the

deflection of a control surface interacting

with the distortions of the lifting surface

where it is mounted. Mass balancing

overcomes flutter at subsonic speeds. The

mass balance requirements for all controls

are that at the neutral point position and at

10o deflection, the product of inertia for the

complete control shall be zero. In addition

the center of gravity must fall within 0.5 Cf

of the hinge line under the most adverse

position where Cf is the mean chord behind

the hinge line of the control surface. When

the mass balance is less, the center of

gravity should always be on, or forward of

the hinge-line. Concentrated mass balance

should not coincide with the nodal line of a

critical structural mode. Location of mass

balance on control linkage and levers is

avoided. At higher compressible speeds,

mass balancing problem becomes difficult

due to increased negative damping. Mass

balance weights should be tested to

substantial dynamic loading and for the

following ultimate attachment strength

conditions:

+1.5 n1g, normal -0.75 n1g,

normal

5.0g laterally 10.0g fore &

aft

Angular acceleration 500 rad/sec2

(about control hinge line)

The backlash in control surface and

systems should be as small as possible. As a

guide the tolerable value range from no

more than .0005 radians for surfaces on

inherently unstable aircraft up to .02 radians

for high lift devices for normal trailing edge

controls at 0.002 radians. Spanwise

distortion of the control surface relative to

the wing, due to air-loads on both the wing

and the control must not be too larger that

physical interference may result. As a guide

the maximum vertical deflection of the


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control relative to the main surface should

not exceed 0.02d, where d is the local

(thickness) of the control surface.

Three general areas make up the field of

aero elasticity testing of scaled models. The

first is where there is no air stream i.e., static

testing to find the stiffness distribution &

vibration testing to find natural frequencies

& mode shapes. The other two areas require

the presence of airstream. Test in second

area involve `steady-state` aeroelastic

phenomenon such as control effectiveness,

whereas third test area includes only

ùnsteady` phenomenon such as flutter and

dynamic stability

References

1, Denis Howe, Àircraft Loading and

Structural layout’, AIAA Educational Series

2004.

2. Howe, D. Àircraft Conceptual

Design Synthesis’, Professional Engineering

Publication Ltd., 2000.

3. Anon., `Part 23, Airworthiness

Standards: Normal Utility, Acrobatics and

Computer Category Airplanes, `Federal

Aviation Regulations, U.S. Government

Printing Office, Washington, DC, Feb. 1991.

4. Anon., `Part 25, Airworthiness

Standards : Transport Category Airplanes,

`Federal Aviation Regulations’, U.S.

Government Printing Office, Washington,

DC, Feb. 1991.

5. Lloyd R. Jenkinson, Paul Simpkin &

Darren Rhodes, `Civil Jet Aircraft Design’ :

AIAA Education Series, 1999.

8. Daniel, P. Raymer, Àircraft Design :

A Conceptual Approach’, AIAA Education

Series, 1992.

6. Darrol Stinton, `Flying Qualities &

Flight Testing of the Aeroplane’, (c) 1996

Blackwell Science Ltd.

7. Ted L. Lomax, `Structural Loads

Analysis for Commercial Transport Aircraft:

Theory & Practice’, AIAA Education Series,

1996.

8. Schuster, D.M., Liu, D.D., and

Huttsell, L.J., ‘Computational Aeroelasticity

: Success, Progress, Challenge,’ Journal of

Aircraft, Vol. 40, No.5, 2003.


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Documents

AIRWORTHINESS: LOADS SPECIFICATIONS · AIRWORTHINESS: LOADS SPECIFICATIONS Air Commodore(retd) S C Gupta Professor & Head Department of Aeronautical Engineering M V J College of Engineering,