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TRANSCRIPT
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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.