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FLIGHTTESTINGLABREPORT

FlightLab

Indian Institute ofTechnology

Kanpur.

Submitted by: (GROUP 1)

Aabhas Srivastava 08AE1024

Praveen Kumar 08AE1012

Raghu V 08AE3006

Vineet Prashant Toppo 08AE1023

C.V. Krishna Koundinya 08AE1004

Department of Aerospace Engineering

Indian Institute of Technology, Kharagpur


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ACKNOWLEDGEMENT

We would like to extend our sincere thanks to Dr. Ajay Mishra for his lucid

lectureson Flight Stability and Control, Mr. Adarsh Mishra for guiding us in our Flight

Lab Experiments. Also, we would like to thank Capt. K. V. Singh, without him we

would never had the opportunityto witness extraordinary maneuvers.

We would also like to extend our thanks to the managementof IIT, Kanpur for

their co-operation throughout the course and also providing us with the facilities.

Also, we would like to thank each and every staff of The FLIGHT LAB for their

Co-operation during theentirecourse.


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CONTENTS

1. Introduction

2. Determination of Centre of Gravity

3. CruiseExperiment

4. Climb Experiment

5. Determination Of SideSlipCoefficient

6. Steady Co-ordinate Turn

7. Dutch Roll Demonstration

8. Phugoid Demonstration

9. Stall Demonstration


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1. INTRODUCTION

PiperSaratogaII

The PA-32 was originally known as the Cherokee Six,deriving from the PA-28

Cherokee series, with heavy modifications.The majordifferences from the PA-

28 were itssix cylinder engine,and six seat configurations.Productiondeveloped

from 1965 through until 1979, with the Cherokee Six-300,Lance, Lance II and

Turbo- Lance.

These were replaced from 1979 by the Saratoga. It was available with fixed

or retracting undercarriage and standard or turbocharged engines. Production

ended in 1985, but in 1995 Piper introduced the Saratoga II HP. The type has

continued to develop in the 1990s, includinga turbocharged version.Thisaircraft

isequipped with a 6-cylinder LycomingIO-540-K1G5engine and a HartzellHC-I3-

YR-1RF propeller.


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The varioussystems of the Piper Saratoga are as follows:

1. Landing Gear

2. Flight Controls

3. Pitot-StaticSystem

4. FuelSystem

5. ElectricalSystem

6. InstrumentPanel

LANDINGGEAR

The airplane is equipped with a retractable tricycle landing gear, which is

hydraulically actuated by an electrically powered reversiblepump. The landinggear is

retracted or extended in about 7 seconds. Emergency Gear extensionsystem allows the

landinggear to freefall, with springassist on the nose gear, into the extended position

where mechanical locksengage. The nose gear issteerable to a 22.5 degree arc each

side of the center through the use of the rudder pedals.The Oleostruts are of the air-

oil type, with normalextensionbeing3.25 .25" for the nose gear and 4.5 .5" for

the maingear under normalstaticload.The standard brake system includestoe brakeson the leftand the right side of rudder pedalsand a hand brake located below the

instruments panel

FLIGHTCONTROLS

Dual flight controls are provided as standard equipment. A cable system provides

actuation of the control surfaceswhen the flight controls are moved in their respective

directions.

The horizontalsurface (stabilizer) featurea trim tab/servo mounted on the trailing

edge. Thistab services the dualfunction of providing trim control and pitch control

forces.

The rudder is conventional design and incorporates a rudder trim. The trim

mechanism isa spring-loaded re-centeringdevice.


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PITOT-STATICSYSTEM

The system suppliesboth pitot and staticpressure for the airspeedindicator,altimeter

and verticalspeed indicator.Pitotand staticpressure are picked up by the pitot head

on the bottom of the left wing. Al alternate static source is provided as standard

equipment.

FUEL SYSTEM

The standard fuel capacity of the Saratoga II HP is107 gallons, of which 102 gallons

are usable.The inboard tank isattached to the wing structure with screws and nut

plates and can be removed for service or inspection.The outboard tanks consist of a

bladder fuel cell that is interconnected with the inboard tank. A flush fuel cap is

located in the outboard tank only. The fuel selectorcontrolhas three positions, one

position corresponding to each wing tank plus an OFF position. A fuel quantity

indicator to measure the fuel not visible through the filler neck in each wing is

installed in the inboard fuel tank. Thisgauge indicatesusable fuel quantities from 5

gallons to 35 gallons in the ground attitude.The sole purpose of this gauge is to assist

the pilot in determining fuel quantities of less than 35 gallons during the preflight

inspection. An electric fuel pump isprovided for use in case of failure of the engine

driven pump.

ELECTRICALSYSTEM

The 14-volt electrical system includes a 12-volt battery for starting and to back up

alternator output. Electricalpower issupplied by a 90-ampere alternator.The battery, a

master switch relay,a voltageregulator and an over voltage relay are locatedbeneath

the floor of the forwardbaggage compartment. Standard electricalaccessoriesinclude

the starter, the electric fuel pump, and the stallwarning horn, the ammeter, and the

annunciatorpanel.The annunciator panellightsare provided only as a warning to the

pilot that a system may not be operating properly, and that the applicable system

gauge should be checked and monitored to determinewhen or if any correctiveaction

is required.


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INSTRUMENTPANEL

The instrument panel is designed to accommodate the customary advanced flight

instrumentsand the normally required power plant instruments.The artificialhorizon

and directionalgyro isvacuum operated and are located in the center of the left-hand

instrument panel. The vacuum gauge is located on the upper left hand instrument

panel. The turn indicator, on the left side, is electrically operated. The radios are

located in the center section of the panel,and the circuit breakers are in the lower right

corner of the panel. An optional radioMASTER switchislocatedon the lower center

instrumentpanel in the switchcluster. It controls the power to allradios through the

aircraftMASTER switch.The radio power switchhas an OFF, and ON position.


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2. DETERMINATIONOFCG

The center-of-gravity (CG) is the point at which an aircraft would balance if it were

possible to suspend it at that point. It isthe mass center of the aircraft, or the theoretical

point at which the entire weight of the aircraftisassumed to be concentrated. Itsdistance

from the referencedatum isdetermined by dividing the totalmoment by the total weight

of the aircraft. The center-of-gravity point affects the stability of the aircraft. To ensure

the aircraftissafe to fly, the center-of-gravity must fall within specifiedlimitsestablished

by the manufacturer.

Center of gravity iscalculatedas follows:

Determinethe weightsand arms of allmass within the aircraft.

Multiply weights by arms for allmass to calculatemoments.

Add the moments of allmass together.

Divide the totalmoment by the total weight of the aircraft to give an overallarm.

The arm that results from this calculationmust be within the arm limits for the center of

gravity that are dictated by the manufacturer. If it is not, weight in the aircraft must

be removed, added (rarely), or redistributed until the center of gravity falls within the

prescribed limits. For the sake of simplicity, center of gravity calculations are usually

performedalong only a single line from the zero point of the referencedatum, usually

the line that represents the roll axis of the aircraft (to calculate fore-aft balance). In

complex situations,more than one line may be separately calculated,e.g., one calculation

for fore- aft balance and one calculation for left-right balance. Weight is calculated

simply by adding up all weight in the aircraft.

This weightmust be within the allowable weight limits for the aircraft. The weightand

moment of fixed portions of the aircraft (engines, wings, etc.) does not change and is

provided by the manufacturer. The manufacturer alsoprovides information facilitating

the calculation of moments for fuel loads. Other removable weight must be properly


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accounted for in the calculation by the operator. In larger aircraft, weightand balanceis

often expressed as a percentage of mean aerodynamicchord, or MAC.

For example, assume that by using the calculation method above, the center of

gravity (CG) was found to be 76 inches aft of the aircraft's datum and the leading

edge of the MAC is 62 inches aft of the datum. Therefore, the CG lies 14 inches aft

of the leading edge of the MAC. If the MAC is 80 inches in length, the percentage of

MAC is found by calculating what percentage 14 is of 80. In this case, one could say

that the CG is 17.5% of MAC. If the allowable limits were 15% to 35%, the aircraft

would be properly loaded.

Calculation

Distance of the referencepoint from NOSE Wheel = 14.2 in

Distance of the referencepoint from REAR Wheel = 109.7 in

94.48 in.

82.92 in.

ForSortie 1:

WEIGHTS L(kg) N(kg) R(kg) TOTAL WEIGHT (kg)

TOW 642 246 656 1544

LW(withoutpassengers)

398 313 405 1116


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Observation

Take-off weight : 1544 kg

Landing weight (excluding wt of passengers and pilot) : 1116 kg

Aspect ratio : 8.167

Plan form area : 16.54 m2

By usingthe C.G formula(above) the center of gravity duringTake-offand Landing are

found to be 94.48" and 82.92" correspondingly to the first sortie.

For Sortie 2:

WEIGHTS L(kg) N(kg) R(kg)TOTAL WEIGHT

(kg)

TOW 643 263 642 1548

LW 395 318 467 1180

Observation

Take-off weight : 1548 kg

Landing we

ight (e

xcluding wt ofpassengers and

pilot) : 1180 kg

By usingthe C.G formula(above) the center of gravity duringTake-offand Landing are

found to be 93.47" and 83.96" correspondingly for the second sortie.


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Incorrectweight and balance in fixed-wingAircraft

When the center of gravity or weight of an aircraft is outside the acceptable range,

the aircraft may not be able to sustain flight, or it may be impossible to maintain the

aircraft in level flight in some or all circumstances. Placing the CG or weight of an

aircraft outsidethe allowedrange can lead to an unavoidable crash of the aircraft.

Incorrectweight and balance in helicopters

The center of gravity iseven more critical for helicoptersthan it is for fixed-wingaircraft

(weight issues remain the same). As with fixed-wing aircraft, a helicopter may beproperly loaded for takeoff, but near the end of a long flight when the fuel tanks are

almostempty, the CG may have shiftedenough for the helicopter to be out of balance

laterally or longitudinally.[2) For helicopters with a single main rotor, the CG is

usually close to the main rotor mast. Improper balance of a helicopter'sloadcan result in

serious controlproblems. In addition to making a helicopter difficult to control, an out-of-

balance loading condition also decreases maneuverability since cyclic control is less

effective in the directionopposite to the CG location.


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The pilot tries to perfectly balancea helicopter so that the fuselage remainshorizontal in

hovering flight, with no cyclic pitch control needed except for wind correction. Since

the fuselageacts as a pendulum suspended from the rotor, changing the center of gravity

changes the angle at which the aircrafthangs from the rotor. When the center of gravity is

directly under the rotor mast, the helicopter hangs horizontal; if the CG istoo far forward

of the mast, the helicopter hangs with itsnose tilteddown; if the CG is too far aft of the

mast, the nose tilts up.

CGforward of forwardlimit

A forward CG may occur when a heavy pilot and passenger take off without baggage or

proper ballastlocatedaft of the rotor mast. Thissituationbecomes worse if the fuel tanks

are locatedaft of the rotor mast because as fuel burns the weightlocatedaft of the rotor

mast becomes less. Thiscondition is recognizable when coming to a hover following a

vertical takeoff. The helicopter will have a nose-low attitude, and the pilot will need

excessive rearward displacement of the cyclic control to maintaina hover in a no-wind

condition. In this condition,the pilot could rapidly run out of rearward cyclic control as

the helicopter consumes fuel. The pilot may also find it impossible to decelerate

sufficiently to bring the helicopter to a stop. In the event of engine failure and the

resultingautorotation,the pilot may not have enough cyclic control to flareproperly for

the landing.

A forward CG will not be as obvious when hovering into a strong wind, since less

rearward cyclic displacement is required as when hovering with no wind. When

determiningwhether a criticalbalanceconditionexists, it isessential to consider the wind

velocity and its relation to the rearward displacement of the cyclic control.


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Weight out of range

Few aircraft imposea minimum weight for flight (althougha minimum pilot weight is

often specified), but all imposea maximum weight. If the maximum weight isexceeded,

the aircraftmay not be able to achieve or sustaincontrolled,level flight. Excessivetake-

off weightmay make it impossible to take off within availablerunway lengths, or it may

completely prevent take-off. Excessive weight in flight may make climbing beyond a

certain altitude difficult or impossible, or it may make it impossible to maintain an

altitude.

CG of aft limit

Without proper ballast in the cockpit, exceeding the aft CG may occurwhen:

A lightweight pilot takes off solo with a full load of fuel locatedaft of the

rotor mast.

A lightweight pilot takes off with maximum baggage allowed in a baggage

compartment locatedaft of the rotor mast.

A lightweight pilot takes off with a combination of baggage and substantial fuel

where both are aft of the rotor mast.

An aft CG conditioncan be recognized by the pilot when coming to a hover following a

vertical takeoff. The helicopter will have a tail-low attitude, and the pilot will need

excessive forward displacement of cyclic control to maintain a hover in a no-wind

condition. If there is a wind, the pilot needs even greater forward cyclic. If flight is

continued in this condition,the pilot may find it impossible to fly in the upper allowable

airspeed range due to inadequate forward cyclic authority to maintain a nose-low

attitude. In addition, with an extreme aft CG, gusty or rough air could accelerate the

helicopter to a speed faster than that produced with full forward cyclic control. In this

case, asymmetry of lift and blade flapping could cause the rotor disc to tilt aft. With

full forward cyclic control already applied, the rotor disc might not be able to be

lowered, resulting in possibleloss of control, or the rotor blades strikingthe tail boom.


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Lateral Balance

In fixed-wing aircraft, lateralbalance is often much lesscritical than fore-aft balance,

simply because most mass in the aircraftislocated very close to itscenter. An exception

is fuel, which may be loaded into the wings, but since fuel loadsare usually symmetrical

about the axis of the aircraft,lateralbalanceis not usually affected.The lateralcenter of

gravity may become important if the fuel is not loadedevenly into tanks on both sides of

the aircraft, or (in the case of small aircraft) when passengers are predominantly on

one side of the aircraft (such as a pilot flying alone in a small aircraft). Small

lateral deviations of CG that are within limits may cause an annoying roll tendency

that pilots must compensate for, but they are not dangerous as long as the CG remains

within limits for the duration of the flight.

For most helicopters, it isusually not necessary to determine the lateral CG for normal

flight instruction and passenger flights. This is because helicopter cabins are relatively

narrow and most optional equipment is located near the center line. However, some

helicopter manuals specify the seat from which solo flight must be conducted. In

addition, if there isan unusualsituation,such as a heavy pilot and a full load of fuel on

one side of the helicopter, which could affect the lateral CG, its position should be

checked against the CG envelope. If carrying external loads in a position that requires

large lateral cyclic control displacement to maintain level flight, fore and aft cyclic

effectivenesscould be dramatically limited.

Fuel dumping and overweightoperations

Many large transport-category aircraftare able to take-offat a greater weight than they

can land.This ispossiblebecause the weight of fuel that the wings can support along

their span in flight, or when parked or taxiing on the ground, isgreater than they can

tolerate during the stress of landingand touchdown, when the support is not distributed

along the span of the wing. Normally the portion of the aircraft's weight that exceeds

the


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Maximum landing weight (but falls within the maximum take-off weight) is entirely

composed of fuel. As the aircraftflies, the fuel burns off, and by the time the aircraftis

ready to land, it isbelow itsmaximum landing weight.However, if an aircraftmust land

early,sometimes the fuel that remainsaboard stillkeeps the aircraft over the maximum

landing weight.When this happens, the aircraftmust either burn off the fuel (by flying in

a holding pattern) or dump it (if the aircraft isequipped to do this) before landing to

avoid damage to the aircraft. In an emergency, an aircraftmay choose to landoverweight,

but this may damage it, and at the very least an overweight landing will mandate a

thorough inspection to check for any damage.

In some cases, an aircraftmay take off overweightdeliberately. An example might be an

aircr

aft being f

erri

edover a very long distance with extra fuel aboard. An overw

eight

take-off typically requires an exceptionally long runway. Overweightoperationsare not

permitted with passengers aboard.


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3. EXPERIMENT:CRUISE

Aim

The static performance characteristics of the Piper Saratoga aircraft which

includesparameters like Zero Lift Drag Coefficient (CDo), Ostwaldefficiency factor (e)

and Induced Drag Co-efficient (k) and variation in control variables like elevator

deflectionand stick force with Lift coefficientare to be determined.

The aim of the cruising experiment is to obtain the curves of power required against

speed for standard weightand sea levelconditions.From these curves, the curve for any

weight and altitude combination can be drawn by suitable scaling. Maximum and

minimum speeds, the speeds for maximum endurance and maximum range can also be

arrived at. In addition, the power available curves provided by the engine performance

charts and the propeller chart make it possible to evaluate the climb performance

characteristicssuch as the maximum rate of climb, the steepest angle of climb, etc. can


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also be found in flight and would serve as a useful basis for comparison with the

theoreticalestimatesmentionedabove.

Instruments used

1. AirspeedIndicator

2. Engine Rpm Indicator

3. ManifoldPressure Gauge

4. Outside Air Temperature Gauge

5. Altimeter

6. Stopwatch.

Procedure

The airplane is made to cruise at four different airspeeds. During the cruise, altitude,

Outside Air Temperature, Engine RPM, Manifold Air Pressure and the timeare noted

from the instrumentpaneland watches (wristwatchand desktop clock). In the same time

stick Force, Angle of attack, Sideslipangle and deflectionangles of aileron, rudder and

elevatorare recorded using the software Lab View in the provided laptop.The recorded

data ispresented in the following t

able:

Altitude

(ft)

Velocity

(Knots)

M.A.P

(inchesof

Hg)

RPM

O.A.T.TimeOn

WW

(hrs)

Timeon

Desktop

(hrs)

1000 80 18 2300 28 1241 1252

1000 85 18 2300 28 1242 1253

1000 90 19 2300 28 1243 1254

1000 95 19 2300 28 1244 1255


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0

0

The above data isfed to the program "CRUISE". The program "CRUISE.exe"calls

various inbuilt subroutineslike"PROPELLER",'BHP'and other standard subroutines to

give the output of (THP)o. The plots of (THP)o*Vo vs. V4

And (THP)o vs. Vo are then

obtained from the output files of CRUISE.Interceptand slope of (THP)o*Vo vs. V4

Are

used to obtainzero lift drag coefficient,Ostwaldefficiency factor(e) and Induced Drag

Co-efficient (k) as follows:

FormulaeUsed

1.

2.

3.

Where,

W isthe TotalTake-offWeight

AR isthe Aspect Ratio.

isthe sea leveldensity (=1.225 kg/m3 )

A plot for estimation of stability characteristics is obtained by using the recorded data

(using LABVIEW) duringflight.


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ResultI ResultII

(ft/s)4 (THP) * Vo o

(hp*ft/s)

Vo

(ft/s)

(THP)o

(hp)

4.16E+08 13368.3 93.630 142.778

5.03E+08 14342.3 95.767 149.762

6.59E+08 16970.4 105.911 160.232

7.82E+08 18038 107.870 167.220

From the above curve,

SlopeIntercept

On Y-Axis

Profile Drag

Coefficient

(Cdo)

Oswald's

Efficiency

(e)

Induced Drag

Coefficient

(K)

1E-05 7813 0.0283 0.475 0.082


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FormulaeUsed

The plots of (dFs/q)/(dCL) and (doe/dCL) are to be obtained for determiningthe neutral

and maneuver points of the airplaneunder stick-fixedand stick-freeconditions,where Fs

is Stick force (in N) taken as the average of the readings recorded at the respective

desktop time in laptop.

CL isthe Coefficient of Lift and isgiven by,

e

isthe elevatordeflection. (in deg)

dW/dT isthe rate of change of weight with respect to timeand isgiven by,


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dw/dt = 0.397 kg/min

FS (N) (deg) Time (min) W (kg) CL FS/q

-7124.97 35.15276 9 1540.427 0.913745 -7.1258

-7043.7 35.02358 10 1540.03 0.830417 -6.40375

-7081.8 35.11045 11 1539.633 0.72534 -5.62516

-7004.28 34.98329 12 1539.236 0.665924 -5.10916

Followingplotsare drawn by usingthe above data given:-


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Results

0.0283

0.082-

0.475

CONCLUSION:

The static performance characteristics of the Piper Saratoga aircraft which

includesparameters like Zero Lift Drag Coefficient (CDo), Ostwaldefficiency factor (e)

and Induced Drag Co-efficient (k) and variation in control variables like elevator

deflectionand stick force with Lift coefficientare determined.


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4. EXPERIMENT:CLIMB

Aim:

The aim of the climb performanceexperimentis to determinethe maximum rates of

climb, and the corresponding speeds at differentaltitudes,and to extrapolate the service

and absolute ceilings for airplane. Minimum timerequired for climb from h1 to h2 can

alsobe evaluated.

Theory:

The governing equations for quasi-steady climb are:


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With the help of the engine

charts and the propeller charts,

the curves of power available

can be plotted againstspeed for

differentaltitudeson the same

coordinate system on which the

power required curves are

plotted. A typicalcurve isshown

in Figureabove.

The speeds at which the two

curves intersectrepresent the minimumand maximum speeds for level flight attainableat

the given altitude. In certain

cases, the stallingspeed may be

greater than the minimum

POWERAVAILABLEANDPOWERREQUIREDCURVES

FORAPROPELLERDRIVENAIRPLANE

speed; in which case this minimumsped losesitssignificance. It can be shown that

under certain restrictions,the speed corresponding to the minimumpower required isthe

speed for maximum endurance and the speed corresponding to the point where the

tangent from the origin touches the power required curve isthe speed for maximum

range. The rate of climb which isgiven by R/C = (Pav - Preq)/w can alsobe evaluatedat

variousspeed and the maximum rate of climb corresponds to the maximum ordinate

between the two

curves.

Instruments used:

1. Airspeedindicator

2. Engine RPM indicator

3. Manifoldpressure gauge

4. Outsideair temperature gauge

5. Altimeter

6. Stopwatch


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Procedure

1. Climb to an altitude h, and set the power to cruise at an airspeed 'V'. Note the

ambient temperature, altitude, engine RPM, and the manifold pressure. Repeat this step

for different speeds. To study the effect of the propeller-pitch, the above trials may

be repeated at the same speeds by setting the propeller to operate at coarse and fine

pitch settings.

2. Establisha steady climb at the maximum continuouspower setting.Record the time

taken to climb through h at variousmean altitudes h, the indicatedairspeed, the engine

rpm, manifoldpressure and ambient air temperature.

3. Calculate the cruise and climb performance characteristics as per the specified

procedure, and present them in the form of charts and plots as shown below:

Velocity

(knots)

H1

(ft)

H2

(ft)

O.A.T M.A.P

(inches

of hg) RPM

Timeon

WW(hrs)

(Time

Differenc

e Betweenh1 and h2)

(sec)

95 500 1000 28 29 2700 1252 36.92

90 500 1000 28 28 2600 1257 35.78

85 500 1000 28 28 2600 1305 30.53

80 500 1000 28 28 2600 1310 34.25


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Plot of Gamma vs. Vo

Plot of Rate of Climb vs. Vo


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Plot of PowerAvailable (THPo Vs Vo) andPower Required (DHPo VsVo)

Results:

From the Plot of Gamma ( ) vs. Vo and Rate of Climb vs. Vo

For Steepest Climb:

4.65o approx

167.39 ft/s

14.22 ft/s

For Max Rate of Climb:

17.2 ft/s

167.67 ft/s

4.035o


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CONCLUSION:

The Plots of Gamma vs. Vo and Rate of Climb vs. Vo did not have anymaxima in the range of velocities where experiments were conducted.


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5. ESTIMATIONOFSIDE-SLIP

COEFFICIENT

Aim:

To estimatethe sideslipcoefficientand find the ratiosbetween the weathercock stability

and aileronyaw stability coefficientsand between C la and dihedraleffect.

Theory:

The conditionsaccompanying the steady sideslipexperimentsare as follows.

Fy = 0; p = 0

r = 0; ay = 0.

Usingthese conditionson the equations for equilibrium in the y-directionduringLateral

DynamicStability Analysis,

On further simplificationsand assumptions, it reduces to:

Further the moment equilibriumequationabout the z directiongives:

= 0, p = 0, r = 0

Now assuming: = 0::we get,


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The moment equilibriumequationabout the x directiongives,

= 0 , p=0,r=0

Now assuming: = 0 , we get,

Procedure:

The airplane is to be flown at four differentbank angles maintaininga fixed velocity,

altitudeand sideslipangle. Duringeach of these 4 bank angles readingswere noted from

the instrumentpanel. Also LabView was used to record the stick force, sideslip angle,

rudder and ailerondeflectionsand the timeat which the 'e'readings are recorded.

For equations (2) and (3), plots of sideslipangleversus rudder, ailerondeflections were

drawn, and their slopesobtained.The sign of these slopesisverified to ensure the lateral

directionalstability of the aircraft. The values of all angles are as given in labview

software and hence are plotted accordingly, except the bank angle which is in

degrees.


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y = 0.067x + 6.100

-25

-24.5

-24

-23.5

-23

-22.5

-22

-21.5

-21

-20.5

-20

-450 -440 -430 -420 -410 -400

Rudder Deflection vs. Sideslip

Rudder

Linear (Rudder)

Beta

RudderDeflection

y = 0.189x - 4.404

-89

-88

-87

-86

-85

-84

-83

-82

-81

-80

-450 -440 -430 -420 -410 -400

Aileron Deflection vs. Sideslip

Aileron

Linear (Aileron)

Beta

Aileron

Deflection


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Bank Angle (deg)/r

/aCy beta (taking beta

in volts as from reading)

1018.51265 5.026227

-0.0849

1518.65328 5.014521

-0.1328

2018.86123 5.009367

-0.1811

2518.99573 4.999852

-0.2265

CONCLUSION:

The various plots required for the analysis of steady slip characteristics have been

obtained and the ratios of required stability coefficients have been computed on the

basis of the given data. It is to be noted that all values obtained from LabView

software are potentiometer readings and hence are in voltage. Since, we have not

been provided with any conversion factor, we have been forced to get our results

using these values only.

y = 0.487x - 92.54

-89

-88

-87

-86

-85

-84

-83

-82

-81

-80

0 10 20 30

Bank Angle vs. Aileron

Aileron

Linear (Aileron)


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6. EXPERIMENT:STEADY

LEVELTURN

Aim:

To determine the lateraland directional control angles required for trim during steady

coordinated turn and to estimatesome lateraland directionalstaticstability derivatives of

the airplane.

Theory:

In a Steady level coordinated turn an airplane turns at a constant altitude and

forward velocity with zero sideslip.The rate of turn iscomputed by usingthe relation:

Where,

r = rate of turn

g = 9 .8 1m/ s^2

= Bank angle


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In addition, we can estimate the stability of the aircraftqualitatively by calculating the

values of 'a/sin'and 'r/r',where a and r are the aileronand rudder deflections

respectively. If 'a/sin'ispositiveand 'r/r' isnegative,we can concludethe aircraft

isstatically stable.

Procedure:

The aircraftwas made to turn through 90 deg at four differentbank angles maintaininga

fixed altitude, speed and zero sideslip. During the turns, various readings were noted

from the instrument panel and LabView was used to record some other data. From

the readings noted down, we can estimatethe rate of turn usingthe formulagiven above.

By applying rolling moment equation,we can obtainthe relationshipshown below,

From the above equation,we can see that the slope of oa Vssin0graph for the three

differentturns should be positive for a statically stable aircraft.Hence, oa vs sin0is

plotted usingthe values of oa recorded by the softwareand 0 valuesnoted in the card.Similarly, by applying yawingmoment equationwe can arriveat the relationshipshown

below,

From the above equation,since Cnr , Cnrare negativewe can expect that the slope of or

Vs r plot should be negative for a statically stable aircraft.Hence, or vs r plot is

constructed usingthe values of or recorded by the software and r valuescalculatedabove.


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Calculation:

Velocity

(knots)

Bank Angle

(deg)

Rate of

turning

(rad/s)

90 10 -22.9818 -84.6223 34.93504 0.041865

90 15 -22.9231 -84.3967 34.84253 0.050361

90 20 -22.9331 -84.4147 34.85433 0.077377

90 25 -22.97 -84.5908 34.91374 0.101865

-84.7915

-84.6223

-84.4530

-84.2838

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45

Sin vs. Aileron Deflection

Sin

sin

Aileron

Deflection(volts)


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CONCLUSION:

The lateraland directional control angles required for trim during steady coordinated

turn and to estimatesome lateraland directionalstaticstability derivatives of the airplane

are determined.


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7.Experiment: DutchRoll Demonstration

Dutch roll isa type of aircraftmotion,consisting of an

out-of-phase combination of "tail-wagging" and

rocking from side to side. This Yaw-roll coupling is

one of the basic flight dynamicsmodes (other include

Phugoid , short period, and spiral divergence). This

motion isnormally welldamped Dutch roll modes can

experiencea degradation in damping airspeeddecrease

and altitude increase. Dutch roll stability can be

artifici

ally incr

eased by the inst

allation of

ayaw

damper. Wings placed well above the center of mass,

sweep back(swept wings) and dihedral wings tends to

increase the roll restoring force, and therefore increase

the Dutch roll tendencies this is why high-winged

aircraft often are slightly anhedral, and transport

category swept wing aircraft are equipped with yaw

dampers.


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8.EXPERIMENT: PHUGOIDEFFECT

The Phugoid is a constant angle of attack but varying pitch angle exchange of airspeed and

altitude. It can be excited by an elevatorsinglet ( a short, sharp deflectionfollowed by a return to

the centered position) resulting in a pitch increase with no change in trim from the cruise

condition. As speed decays, the nose will drop below the horizon. Speed will increase, and

the nose will climb above the horizon. Periods can vary from under 30 seconds for lightaircraft to minute for larger aircraft. Micro light aircraft typically show a Phugoid periods of

15-25 seconds, and it has been suggested that birds and model airplane shown convergence

between the Phugoid and short periodmodes.


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10. EXPERIMENT:STALL

Flow separationbegins to occur at smallangle of attack whileattached flow over the

wing isstill dominated.Asangle of attack increases,the separated regions on the top of

the wing increase in sizeand hinder the wing'sability to create lift. At the criticalangle

of attack, separated flow isso dominantthat further increases in angle of attack produces

less lift and vastly more drag. (Note, airflowdoesn't really separate from the wing, a

vacuum does not magically emerge there.

Rather, cleanlaminar flow gets pulledaway by messy turbulent flow " Flow Separation"isa usefulabstraction though)


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11. ConclusionThe flight lab experiments were successfully conducted. The experiments were

performed in two sorties each extending for a period of approx 50-60 min. The analysis of

the Climb and Cruise experiments were done successfully. The performance values

obtained were in close proximity with the expected results within the error limit. In steady

state slip and steady turn experiments, the ratios of the stability coefficients were obtained

which were also within the error limit. Apart from these a demonstration flight was also

organized where we experienced the different modes of flight like Dutch Roll and the

Phugoid.

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