FVD Lab Manual

8/13/2019 FVD Lab Manual

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MLR Institute of Technology Laxman Reddy Avenue, Dundigal Police Station Road,Gandimysamma XRoad, Quthbullapur (M), R.R. Dist - 500 043.Ph: 08418 204066, 204088www.mlrinstitutions.ac.in Email: [email protected]

LAB MANUAL

Subject : F light Vehi cle Design

Academic Year : 2013-2014

Branch & Year : AERO-A II I Year I I SEM

Name of the Faculty : B.SRIKANTH & S.RAVI KANTH

Department : Aeronautical Engineer ing
http://www.mlrinstitutions.ac.in/http://www.mlrinstitutions.ac.in/mailto:[email protected]:[email protected]:[email protected]:[email protected]://www.mlrinstitutions.ac.in/


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LIST OF EXPERIMENTS

INTRODUCTION

PART I

AIRCRAFT DESIGN AND WEIGHT ESTIMATION NOMENCLATURE

Experiment1: Aircraft conceptual sketch and its gross weight estimation algorithm

Experiment2: Preliminary weight estimation (Rubber sizing)

Experiment3: Trade off study on initial (Rubber sizing)

Experiment4: Fixed sizing

Experiment5: Load or Induced Drag Estimation

Experiment6: Estimate the Critical Mach number for an Airfoil

PART II

AIRCRAFT PERFORMANCE

Experiment: 7 Static Performance: Thrust required curve

Experiment: 8 Static Performance: Power required curve


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INTRODUCTION

CLASSIFICATION OF AIRCRAFT BASED ON TYPE, ROLE

AND MISSION

Aim:

To study the classification and designation of aircraft based on type, role and mission and

corresponding mission profile.

Procedure:

Three Views of a Commercial & Military Aircraft

Military Aircraft


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Commercial Aircraft

Classification based on type:

An aircraft is a machine that is able to fly by gaining support from the air, or, ingeneral, the atmosphere of a planet. It counters the force of gravity by using either static lift

or by using the dynamic lift of an airfoil, or in a few cases the downward thrust from jet

engines.

The human activity that surrounds aircraft is called aviation. Crewed aircraft are

flown by an onboard pilot, but unmanned aerial vehicles may be remotely controlled or self-

controlled by onboard computers. Aircraft may be classified by different criteria, such as lift

type, propulsion, usage, and others.

1. Lighter than air - Airships, Aerostats

Aerostats use buoyancy to float in the air in much the same way that ships float on the

water. They are characterized by one or more large gasbags or canopies, filled with a

relatively low-density gas such as helium, hydrogen, or hot air, which is less dense than the

surrounding air. When the weight of this is added to the weight of the aircraft structure, it

adds up to the same weight as the air that the craft displaces.


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2. Heavier than air aerodynes

Heavier-than-air aircraft must find some way to push air or gas downwards, so that a

reaction occurs (by Newton's laws of motion) to push the aircraft upwards. This dynamic

movement through the air is the origin of the term aerodyne. There are two ways to produce

dynamic up thrust aerodynamic lift, and powered lift in the form of engine thrust.

a. Power driven

I. Airplane

i. Land Plane

ii. Sea Plane

iii. Amphibian

II. Rotorcraft

i. Helicopter

ii. Gyrocopter

iii. Autogyro

III. Ornithopter

b. Non Power driven

I. Gliders

II. Sail Plane

Unpowered aircraft

Gliders are heavier-than-air aircraft that do not employ propulsion once airborne.

Take-off may be by launching forward and downward from a high location, or by pulling into

the air on a tow-line, either by a ground-based winch or vehicle, or by a powered "tug"

aircraft. For a glider to maintain its forward air speed and lift, it must descend in relation to

the air (but not necessarily in relation to the ground). Many gliders can 'soar' gain height

from updrafts such as thermal currents. The first practical, controllable example was designed

and built by the British scientist and pioneer George Cayley, whom many recognise as the

first aeronautical engineer.

Powered Aircraft

Propeller aircraft use one or more propellers (airscrews) to create thrust in a forward

direction. The propeller is usually mounted in front of the power source in tractor

configuration but can be mounted behind in pusher configuration. Variations of propeller

layout include contra-rotating propellers and ducted fans.


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Many kinds of power plant have been used to drive propellers. Early airships used

man power or steam engines. The more practical internal combustion piston engine was used

for virtually all fixed-wing aircraft until World War II and is still used in many smaller

aircraft. Some types use turbine engines to drive a propeller in the form of a turboprop or

propfan. Human-powered flight has been achieved, but has not become a practical means of

transport. Unmanned aircraft and models have also used power sources such as electric

motors and rubber bands.

Classification Based on Role:

I. Civil

a. Commercial

i) Passenger

ii) Cargo

b. General Aviation

1) Aerial Burial

2) Aerial Photography

3) Aerobatics

4) Air Ambulance

5) Air Charter

6) Air Shows

7) Rush flying

8) Non commercial air cargo flight

9) Crop dusting

10) Emergency management

11) Flight Trainer

12) Forest fire fighting

13) Gliding

14) Parachuting

15) Personal Transportation

16) Air Police and Patrol

17) Resource exploration

18) Tourism

II. UAV

1) Micro UAV2) Close Range UAV [CR-UAV]


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3) Medium Range UAV [M-UAV]

4) Combat UAV [C-UAV]

III. MILITARY AIRCRAFT

1 Combat aircraft

i. Fighter

ii. Bomber

iii. Attack aircraft

iv. Electronic warfare aircraft

v. Maritime patrol aircraft

vi. Multirole combat aircraft

2 Non-combat aircraft

i. Military transport aircraft

ii. Airborne early warning and control

iii. Reconnaissance and surveillance aircraft

iv. Experimental Aircraft

Military Aircraft the designation system

The Russian aircraft designation system is based on the company name which

manufactured it

Eg: Mikoyan-Gurevich [MiG]

Sukhoi [Su]

Tupolev [Tu]

Yakovlev [yak]

Ilyushin [II]

Antonov [An]

A U.S. military aerospace vehicle designation is also known as an "MDS

Designation". MDS stands for "Mission-Design-Series", naming the three most important

components of the designation

A Attack B - Bomber

C - Cargo/Transport


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D - Drone Director E - Special Electronics Installation F Fighter

G - Glider H - Search and Rescue (SAR) I - Interceptor J - Special Test (temporary) K - Tanker L - Polar M Multipurpose

N - Special test (permanent) O Observation P- Patrol Q - Drone R - Reconnaissance S - Anti-submarine T - Trainer

U - Utility V - Staff/VIP W - Weather X- Experimental Y- Prototype Z- lighter than air

Combat AircraftCombat aircraft, or "Warplanes", are divide broadly into multi-role, fighters, bombers

and attackers, with several variations between them, including fighter-bombers, such as the

MiG-23, ground-attack aircraft, such as the Soviet Ilyushin Il-2 Shturmovik Also included

among combat aircraft are long-range maritime patrol aircraft, such as the Hawker Siddeley

Nimrod and the S-3 Viking that are often equipped to attack with anti-ship missiles and anti-

submarine weapons.


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Fighter

The main role of fighters is destroying enemy aircraft in air-to-air combat, offensive

or defensive. Many are fast and highly maneuverable. Escorting bombers or other aircraft is

also a common task. They are capable of carrying a variety of weapons, including machine

guns, cannons, rockets and guided missiles. Many modern fighters can attack enemy fighters

from a great distance, before the enemy even sees them. Examples of air superiority fighters

include the F-22 Raptor and the MiG-29. WWII fighters include the Spitfire, the P-51

Mustang and Bf 109. An example of an interceptor (a fighter designed to take-off and quickly

intercept and shoot down enemy planes) would be the MiG-25. An example of a heavy

fighter is the Messerschmitt Bf 110. The term "fighter" is also sometimes applied to aircraft

that have virtually no air-air capability for example the A-10 ground-attack aircraft is

operated by USAF "Fighter" squadrons.

Bomber

Bombers are normally larger, heavier, and less maneuverable than fighter aircraft.

They are capable of carrying large payloads of bombs. Bombers are used almost exclusively

for ground attacks and not fast or agile enough to take on enemy fighters head-to-head. A few

have a single engine and require one pilot to operate and others have two or more engines and

require crews of two or more. A limited number of bombers, such as the B-2 Spirit, have

stealth capabilities that keep them from being detected by enemy radar. An example of a

conventional modern bomber would be the B-52 Stratofortress. An example of a WWII

bomber would be a B-17 Flying Fortress. Bombers include light bombers, medium bombers,

heavy bombers, dive bombers, and torpedo bombers. The U.S. Navy and Marines have

traditionally referred to their light and medium bombers as "attack aircraft".

Attack aircraft

Attack aircraft can be used to provide support for friendly ground troops. Some are

able to carry conventional or nuclear weapons far behind enemy lines to strike priority

ground targets. Attack helicopters attack enemy armor and provide close air support for

ground troops. An example historical ground-attack aircraft is the Soviet Ilyushin Il-2

Shturmovik. Several types of transport airplanes have been armed with sideways firing

weapons as gunships for ground attack. These include the AC-47 and AC-130 aircraft.

Electronic warfare aircraft

An electronic warfare aircraft is a military aircraft equipped for electronic warfare

(EW) - i.e. degrading the effectiveness of enemy radar and radio systems.


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Maritime patrol aircraft

A maritime patrol aircraft fixed-wing military aircraft designed to operate for long

durations over water in maritime patrol roles in particular anti-submarine, anti-ship and

search and rescue.

Multirole combat aircraft

Many combat aircraft today have a multirole ability. Normally only applying to fixed-

wing aircraft, this term signifies that the plane in question can be a fighter or a bomber,

depending on what the mission calls for. An example of a multirole design is the F/A-18

Hornet. A WWII example would be the P-38 Lightning.

Some fighter aircraft, such as the F-4, are mostly used as 'bomb trucks', despite being

designed for aerial combat

Non-combat aircraft

Non-combat roles of military aircraft include search and rescue, reconnaissance,

observation/surveillance, Airborne Early Warning and Control, transport, training, and aerial

refueling.

Many civil aircraft, both fixed wing and rotary wing, have been produced in separate

models for military use, such as the civilian Douglas DC-3 airliner, which became the

military C-47 Skytrain, and British "Dakota" transport planes, and decades later, the USAF's

AC-47 aerial gunships. Even the fabric-covered two-seat Piper J3 Cub had a military version.

Gliders and balloons have also been used as military aircraft; for example, balloons were

used for observation during the American Civil War and during World War I, and military

gliders were used during World War II to deliver ground troops in airborne assaults.

Military transport aircraft

Military transport (logistics) aircraft are primarily used to transport troops and war

supplies. Cargo can be attached to pallets, which are easily loaded, secured for flight, and

quickly unloaded for delivery. Cargo also may be discharged from flying aircraft on

parachutes, eliminating the need for landing. Also included in this category are aerial tankers;

these planes can refuel other aircraft while in flight. An example of a transport aircraft is the

C-17 Globemaster III. A WWII example would be the C-47. An example of a tanker craft

would be the KC-135 Stratotanker. Helicopters and gliders can transport troops and supplies

to areas where other aircraft would be unable to land.

Calling a military aircraft a "cargo plane" is incorrect, because military transport

planes also carry paratroopers and other soldiers.


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Airborne early warning and control

An airborne early warning and control (AEW&C) system is an airborne radar system

designed to detect aircraft, ships and vehicles at long ranges and control and command the

battle space in an air engagement by directing fighter and attack aircraft strikes. AEW&C

units are also used to carry out surveillance, including over ground targets and frequently

perform C2BM (command and control, battle management) functions similar to an Airport

Traffic Controller given military command over other forces. Used at a high altitude, the

radars on the aircraft allow the operators to distinguish between friendly and hostile aircraft

hundreds of miles away.

Reconnaissance and surveillance aircraft:

Reconnaissance aircraft are primarily used to gather intelligence. They are equipped

with cameras and other sensors. These aircraft may be specially designed or may be modified

from a basic fighter or bomber type. This role is increasingly being filled by satellites and

unmanned aerial vehicles (UAVs).

Surveillance and observation aircraft use radar and other sensors for battlefield

surveillance, airspace surveillance, maritime patrol and artillery spotting. They include

modified civil aircraft designs, moored balloons and UAVs.

Experimental Aircraft

Experimental aircraft are designed in order to test advanced aerodynamic, structural,

avionic, or propulsion concepts. These are usually well instrumented, with performance data

telemetered on radio-frequency data links to ground stations located at the test ranges where

they are flown. An example of an experimental aircraft is the XB-70 Valkyrie.

Result:

All study and classification and designation of aircraft based on type role and mission is

completed.


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PART - I

AIRCRAFT DESIGN AND WEIGHT ESTIMATIONNOMENCLATURE

Weight components of airplane explained as follows:

1) Crew weight ( cW ):

The crew comprises the people necessary to operate the airplane in flight.

e.g., Pilot, Co-pilot, Airhostess etc.

2) Payload weight ( pW ):

The payload is what the airplane is mentioned to transport passengers, baggage,

freight etc. (Military use the payload includes bombs, rockets and other disposable ordnance).

3) Fuel weight ( f W ):

This is the weight of the fuel in the fuel tanks. Since fuel is consumed during the

course of flight. f W is a variable, decreasing with time during the flight.

4) Empty weight ( eW ):

This is weight of everything else-the structure engines (with all accessory equipment),

electronic equipment landing gear, fixed equipment and anything else that is not crew,

payload or fuel.

5) Gross weight ( 0W ):

The sum of these weights is the total weight of the airplane 0W . Gross weight or total

weight 0W varies through the flight because fuel is being consumed. The design take off

gross weight 0W is the weight of the airplane at the instant it begins its mission. It includes

the weight of the fuel.

e f pc W W W W W 0

W W W

W W W

W W W e f pc0

00

0

00

0

1

)(

W W

W

W W W

W e f

pc ---------------------------------- (1)


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Estimation of empty weight fraction ( 0W W e ):

The empty weight fraction ( 0W W e ) can be estimated from data based on

a) Historical data and tables

b) Refined sizing data and tables

Estimation of fuel fraction ( 0W W f ):

The aircrafts fuel supply is available for performing the mission. The other fuel

includes reserve fuel, trapped fuel (which is the fuel which cannot be pumped out of the

tanks).

Fuel fraction ( 0W W f ) is approximately independently of aircraft weight. Fuel

fraction will be estimated based on the mission to be flown.

Mission profiles:

Typical mission profiles for various types of aircraft are shown in Fig1. The simple

cruise mission is used for many transport and general aviation designs, including home built.

Following are the briefly explained the terms that are used in mission profiles:


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Warm Up and Take-Off :

Warm Up is the engine start up for the airplane kept idling for some time to warm up.

Take Off is the point where aircraft is made lift off from ground. It is the motion after

warm up i.e., moving of airplane after starting and till it lifts off from the ground.

Climb:

It is between take-off (TO) and cruise (stead level flight with constant speed) Increase

in height until airplane achieves steady level flight.

Cruise:

It is the steady level flight to cover the mission distance. The mission distance is

called Range.

Loiter:

Represent the airplane spending in air for some fixed number of minutes near airport

before getting the clearance from airport signal or simple spending some time to

collect data of some mission (Terrain data).

Dash:

It is the mission that must be flown at just a few hundred numbers of feet of the

ground for low level strike.

Landing:

It is the aircraft landing on the runway till stopping of engine.

Estimation of mission segment weight fractions:

The various mission segments (legs) are numbered starting from zero denoting, the

start of the mission. Mission leg one is usually engine warm up and take-off. The remaining

legs are sequentially numbered. For example in the simple cruise mission the legs could be

numbered as (0) warm-up and take-off, (1) climb (2) cruise (3) loiter and (4) landing.

Similarly, the aircraft weight at end of each mission is denoted by iW . Denoting i-th

segment as mission segment weight

0W =Beginning airplane weight (Take off gross weight)

1W =Weight of the airplane at end of warm-up and take-off

2W =Weight of the airplane at end of climb.

3W =Weight of the airplane at end of cruise

4W =Weight of the airplane at end of loiter.


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5W =Weight of the airplane at end of landing.

4

5

3

4

2

3

1

2

0

1

0

50 ... W

W W W

W W

W W

W W

W W

W W x

So in general it can be written as

13

4

2

3

1

2

0

1

00 ...

i

ii x W

W W W

W W

W W

W W

W W

W W

Warm-u p/take-off , cli mb and landing weight f ractions:

The warm-up, take-off and landing weight fractions can be estimated historically from

Table2.

Specif ic fuel consumpti on (C):

It is the rate of fuel consumption divided by the resulting thrust. Typical values are

depicted in Table3 and Table4 for jet and propeller aircrafts respectively. If the aircraft is propeller, then C should be replaced by )550( pbhpV C C

Crui se segment weight f r action:

Weight fraction for cruise segment is found using Breguet range formula.

i

iW

W D L

C V

R 1ln R = range, C = specific fuel consumption


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D LV RC

W W

i

i exp1

V = velocity, L/D = lift to drag ratio

L oiter segment weight fr action:

Weight fraction for loiter segment is found using Endurance formula.

i

i

W W

C D L

E 1ln E = endurance or loiter time, C = specific fuel consumption

D L

EC W W

i

i exp1

V = velocity, L/D = lift to drag ratio

The most efficient cruise is velocity for propeller aircraft occurs at velocity yielding max

L/D, where as for the most efficient cruise for a jet aircraft occurs at slightly at a higher

velocity yielding an L/D of 86.6% of the maximum L/D

Type of aircraft Cruise Loiter

jet 0.866 (L/D)max L/D max

propeller L/D max 0.866 (L/D) max

For any mission segment i the mission segment weight fraction is expressed as

1ii W W . xW (Assuming x segments are present for total mission profile) is the aircraft

weight at end of the mission. 0W W x ratio can be used to calculate fuel fraction.

)(1 00 W W W W x f

At the end of the mission, the fuel tanks are not completed empty, typically a 6% allowance

is made for reserve and trapped fuel

)/(106.1 00 W W W W x f

Estimate of gross weight at take-off ( 0W ):

0W W e is function of 0W , 0W W f is also a function of 0W . 0W is calculated from

equation(1) through process of iteration. 0W is taken a guess value and, then RHS value of

equation(1) is calculated which should match the value of assumed, if it doesnt, increment

the assume by some value and iterate it. This process is continued till the absolute difference

of RHS value and assumed value is the least and that iteration step will be your nearest

solution.


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EXPERIMENT 1

AIRCRAFT CONCEPTUAL SKETCH AND ITS GROSS

WEIGHT ESTIMATION ALGORITHMAIM:

Write the request for proposal for the particular aircraft, draw the conceptual sketch of

the aircraft for given type of aircraft, draw the mission profile and write generic algorithm for

gross take-off weight estimation

THEORY:

CONCEPTUAL DESIGN:

Conceptual design begins with a specific set of design requirements established fromcustomer or a company-generated guess what future customers may need.

Design requirements include:

a) Aircraft range

b) Payload

c) Take-off distance

d) Landing distance

e)

Maneuverability and speed requirementsDesign begins with innovative idea rather than as a response to a given requirement.

Before design a decision is made to what technologies to incorporate, it must use only

currently available technologies as well as existing engines and avionics. If designed to build

in more distant future, then an estimate technological state of the art must be made to

determine which emerging technologies will be ready for use at that time.

Design begins drawing with a conceptual sketch like shown in Fig1. Good conceptual

sketches start with approximate sketch of following:


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1) Wing

2) Tail geometries

3) The fuselage shape

4) The internal locations of the major components such as the:

a) Engine

b) Cockpit

c) Payload/passenger compartment

d) Landing gear

e) Fuel tanks.

SIZING :

The conceptual sketch is used to estimate aerodynamics and weight fractions by comparisons

to previous designs. These estimates are used to make a first estimate of the required total

weight and fuel weight to perform the design mission.

First order sizing provides the information to needed to develop an initial design layout in

three view format. This three view drawing is completed with the internal arrangement in

detail. The initial layout is analyzed to determine if it will perform the mission as indicated by the first-order sizing.

ALGORITHM FOR GROSS TAKE-OFF WEIGHT ESTIMATION:

Following steps are involved in gross take-off weight estimation:

1) Study the design objectives.

2) Sizing mission starts here.

3) Aspect ratio selection is done here.

4) Sketch the layout in three views.5) Select L/D ratio and engine specific fuel consumption.


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6) Estimate fuel weight fraction.

7) Select empty weight fraction (Historical trends).

8) Guess initial gross weight.

9) Calculate gross weight from equation.

10) Iterate for gross weight by going to step8, until guess and calculated are matched.

The following flow chart explains the same algorithm as explained previous

PROCEDURE:

1. Write the request for proposal for the given aircraft. It should be in the form of

parameters and requirements for the aircraft.

2. Draw the conceptual sketch of the aircraft as explained in theory.

3. Draw the mission profile for the aircraft.

4. What do you understand by flight vehicle design? Explain it with various examples.

5. What do you understand by weight estimation and write the algorithm for gross take-off weight estimation.

RESULT:

The take-off weight can be estimated by doing the iterations, until we get,

W 0 guess = W 0 Calculated


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EXPERIMENT 2

PRELIMINARY WEIGHT ESTIMATION (RUBBER SIZING)

AIM

To estimate take-off gross weight for the given aircraft and its mission profile using weight

estimation algorithm.

THEORY

Study the theory Part1: Aircraft design and weight estimation nomenclature on

page1 for basics that is needed for the experiment.

Given empty weight fractions from historical trends (preliminary design): The empty weight

fraction can be estimated from Table1 based on the aircraft type and wing sweep.

Given warm-up/take-off, climb and landing weight fractions from historical trends: The

warm-up, take-off and landing weight fractions can be estimated historically from Table2.


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Requirements:

Aircraft type, engine type, wing sweep type, mission profile, crew weight, payload

weight, specific fuel consumption, L/D ratio.

Procedure:

Estimation of gross weight, calculated using following steps:

1. Calculate the mission weight fraction of individual segment:

1) Take-off ( 01 W W ): This is taken from Table2.

2) Climb ( 12 W W ): This is taken from Table2.

3) Landing ( 45 W W ): This is taken from Table2.

4) Cruise:Weight fraction for cruise segment is found using Breguet range formula.

i

i

W W

D L

C V

R 1ln

D LV

RC W W

i

i exp1

Where R = range,

C = specific fuel consumption

V = velocity,

L/D = lift to drag ratio

5) Loiter


i

i

W W

C D L

E 1ln

D L

EC W W

i

i exp1

Where E = endurance or loiter time,

C = specific fuel consumption,

V = velocity,

L/D = lift to drag ratio

6) Empty Weight fraction: The empty weight fraction can be estimated from Table1

based on the aircraft type and wing sweep.


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2. Calculate gross weight of the aircraft from following equation which is function of 0W .

00

0

1

)(

W W

W W

W W W

e f

pc (1)


equation(1) through process of iteration. 0W has to be assumed, then RHS value of




solution. This is done using following iteration table.

Iteration.

Guess

weightEmpty

weight

Fuel

weight

Calculated

weight

Difference =

guess-calculated

3) Plot graph for calculated weight, guess weight versus iteration number from above table

results and compare them in a single graph.

RESULT:

The iterations should be done until the difference is zero .


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EXPERIMENT 3

TRADE OFF STUDY ON INITIAL SIZING (RUBBER SIZING)

AIM:

To study trade off on initial sizing (weight estimation) by taking following

parameters.

a) Range trade off,

b) Payload trade off.

THEORY:

Study the theory Part1: Aircraft design and weight estimation on page1 for basics,

needed for the experiment.

Given Empty weight fractions from historical trends (preliminary design): The empty

weight fraction can be estimated from Table1 based on the aircraft type and wing sweep type.

Given warm-up/take-off, climb and landing weight fractions from historical trends:The warm-up, take-off and landing weight fractions can be estimated historically from

Table2.


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REQUIREMENTS:

Aircraft type, engine type, wing sweep type, mission profile, crew weight, specific

fuel consumption, L/D ratio, different range values, different payload weight values.

PROCEDURE:I) Estimation of gross weight for selected range value through calculations using following

steps:





4) Cruise:


i

i

W W

D L

C V

R 1ln

D LV

RC W W

i

i exp1

Where R = range, C = specific fuel consumption

V = velocity, L/D = lift to drag ratio5) Loiter


i

i

W W

C D L

E 1ln

D L

EC W W

i

i exp1

Where E = endurance or loiter time, C = specific fuel

consumption, V = velocity, L/D = lift to drag ratio6) Empty Weight fraction: The empty weight fraction can be estimated from Table1

based on the aircraft type and wing sweep.

2. Calculate gross weight of the aircraft from following equation which is function of 0W

00

0

1

)(

W W

W

W W W

W e f

pc (1)


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Iteration. No

Guess

weight

Empty

weight

fraction

Fuel weight

fraction

Calculated

weight

Difference=

guess-cal

3) Plot graph between guess gross weight, calculated gross weight versus iteration number.

4) Steps 1 to 3 is repeated for second range, third range

5) Plot a graph between calculated gross weight versus range selected.

II) Calculate gross weight for different selected payload weight values using step1 to 3 and

plot a graph between calculated gross weight and payload weight values.

RESULT:

The iterations should be done until the difference between the guess and calculated

weights is equal to zero.


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EXPERIMENT 4

FIXED ENGINE SIZING

AIM

To estimate take-off gross weight for the given aircraft and its mission profile using weight

estimation algorithm and calculate the effect of fixed sizing on range.

THEORY

Study the theory Part1: Aircraft design and weight estimation nomenclature on

page1 for basics, needed for the experiment.

Given empty weight fractions from historical trends (preliminary design): The empty weight

fraction can be estimated from Table1 based on the aircraft type and wing sweep.

Given warm-up/take-off, climb and landing weight fractions from historical trends: The

warm-up, take-off and landing weight fractions can be estimated historically from Table2.


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REQUIREMENTS:

Aircraft type, engine type, wing sweep type, mission profile, crew weight, payload

weight, specific fuel consumption, L/D ratio.

PROCEDURE:

Estimation of gross weight, calculated using following steps:





4) Cruise:


i

i

W W

D L

C V

R 1ln

D LV

RC W W

i

i exp1

Where R = range, C = specific fuel consumption

V = velocity, L/D = lift to drag ratio5) Loiter


i

i

W W

C D L

E 1ln

D L

EC W W

i

i exp1

Where E = endurance or loiter time, C = specific fuel

consumption, V = velocity, L/D = lift to drag ratio

EMPTY WEIGHT FRACTION :

The empty weight fraction can be estimated from Table1 based on the aircraft

type and wing sweep.

2. Calculate gross weight of the aircraft from below equation which is function of 0W

00

0

1

)(

W W

W W

W W W

e f

pc (1)


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Iteration. No

Guess

weight

Empty

weight

fraction

Fuel weight

fraction

Calculated

weight

Difference=

guess-cal

3) Once the gross weight is obtained, fixed sizing is done by keeping guess gross weight,

empty weight fraction as constant values, and vary(increment or decrement) fuel weight

fraction.

Iteration.

No

Guess

weight

(constant)

Empty

weight

fraction(constant)

Fuel weight

fraction(vary

this)

Calculated

weight

Difference=

guess-cal

RESULT:

The iterations should be done until the difference between the guess and calculated

weights is equal to zero.


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EXPERIMENT 5

LOAD OR INDUCED DRAG ESTIMATION

AIM

To find drag due to lift or Induced drag for the following aircrafts using Oswalds

span efficiency method and Leading edge suction method.

1) Straight wing aircraft,

2) Swept wing aircraft,

3) Supersonic aircraft.

REQUIREMENTS Aspect ratio Coefficient of lift Sweep of leading edge Speed of aircraft.

THEORY

The induced drag coefficient of moderate angle of attack is proportional to square of

the lift coefficient with a proportionality factor called the drag -due-to-lift- factor or K 2

L D KC C (1)

Following are the two methods to estimate drag -due-to-lift- factor or K :

1) Oswalds span efficiency method

2) Leading edge suction method

1) Oswalds span efficiency method:

According to classical wing theory, the induced drag coefficient of 3D-Wing with an

elliptical lift distribution equals the square of lift coefficient divided by A (A = Aspect

Ratio or Effective Aspect Ratio)

K = Ae 1

(2)

A = Aspect Ratio

effective A = Effective Aspect Ratio

e = Oswalds spa n efficiency (The value of e varies from 0.7 to 0.85)

Effective Aspect Ratio for

End-plates: )/9.11( bh A Aeffective (3)


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h = height of Endplate

Winglets: A Aeffective 2.1 (4)

Straight wing aircraft:64.0)045.01(78.1 68.0 Ae (5)

Swept wing aircraft:

1.3))(cos045.01(61.4 15.068.0 LE Ae (6)

Supersonic aircraft:

LE M A

M Ae cos

2)1(4

]1[2

2

(7)

LE = Sweep angle of leading edge

Disadvantages of Oswald span efficiency method:

1) Ignores the variation of K with lift coefficient.

2) This doesnt include the effects of the change in viscous separation as lift

coefficient is changed.

Wing Details


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2) Leading Edge Suction Method:

This is a semi-empirical for estimation of K allows for the variation of K with lift

coefficient and Mach number. Due to the rapid curvature at the leading edge, there creates a

pressure drop on the upper part of the leading edge. The reduced pressure exerts a suction

force on the leading edge in a forward direction. This leading edge suction force S is in the

direction perpendicular to the normal force N.

1000 )1( K S SK K (8)

Where, K = drag-due-to-lift-factor

A

K

1100

l C K

10

l C = slope of the lift curve, angle taken in radians

S = Leading edge suction factor


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PROCEDURE

1. Select a value of aspect ratio and calculate the Ostwald efficiency factor e using any

one of the equations (5), (6) or (7) based on aircraft type.

2. Calculate drag-due-to-lift-factor or K using equation(2).

3. Calculate coefficient of drag DC (induced drag) using equation (1).

4. Select a value of coefficient of lift LC for corresponding aspect ratio.

5. Iterate step(1) through (3) for varying/incrementing aspect ratio and coefficient of lift

LC .

6. Plot graph between LC verses DC and LC versus K.

RESULT:

Hence, the drag due to lift or Induced drag for the following aircrafts using Oswalds

span efficiency method and Leading edge suction method was found.


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EXPERIMENT 6

ESTIMATE THE CRITICAL MACH NUMBER FOR AN

AIRFOIL

AIM 1. Estimate the Critical Mach number for given airfoil at the specified angle of attack.

Two approaches to estimate critical Mach number

a) Graphical method

b) Analytic method

2. Compare the results obtained in each method.

THEORY :

Critical Mach number: Free stream Mach number at which sonic flow is first obtained

somewhere on the surface of the airfoil is called the Critical Mach number of the airfoil.

11)1(22 )1(2

2,

M M

C cr p (1)

2

0,

1 M

C C p p (2)

11)1(22

1

)1(2

22

0,

cr

cr cr

p M M M

C (3)

PROCEDURE

a) Graphical Method:1) The Graphical method involves following steps:

2) Obtain a plot of Cp versus M from equation (1). This is illustrated by curve A

in Fig1. The curve is a fixed universal curve that is used for all.


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Fig1. Determination of Critical Mach number.

3) For low-speed, incompressible flow, obtain the value of the minimum pressure

coefficient corresponds to the point of maximum velocity on the airfoil

surface. This minimum value of 0, pC must be given to you (from experimental

measurement or theory). This 0, pC is shown as point B in Fig1

4) From the equation (2), plot the variation of this coefficient versus M . This isillustrated by curve C in Fig1.

5) Where curve C intersects curve A, the minimum pressure coefficient on the

surface of the airfoil is equal to the critical pressure coefficient. This

intersection point is denoted by point D in Fig1. For the conditions associated

with this point, the maximum velocity on the airfoil surface is exactly sonic.

The value of M at point D is then by definition, the Critical Mach number.

b) Analytic Method:

Equation (2) gives the variation of pC at a given point on the airfoil surface as a

function of M . At some location on the airfoil surface 0, pC will be a minimum value,

corresponding to the point of maximum velocity on the surface. The value of the minimum

pressure coefficient will increase in absolute magnitude as M is increased owing to the

compressibility effect.


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Hence equation(2) with 0, pC being minimum value on the surface of the airfoil at

essentially incompressible flow conditions ( M < 0.3) gives the value of minimum pressure

coefficient at a higher Mach number M at some value of M the flow velocity will become

sonic at the point of minimum pressure coefficient. The value of the pressure coefficient at

sonic conditions is the critical pressure coefficient, given by equation (1). When the flow

becomes sonic at the point of minimum pressure, the pressure coefficient given by equation

(2) is the value given by equation (1). Equating these two relations we have equation (3).

The value of M that satisfies equation (3) is the value when the flow becomes sonic

at the point of maximum velocity (minimum pressure). That is the value of M obtained

from equation (3) is the critical Mach number for the airfoil.

Equation (3) must be solved implicitly for M by trial and error, guessing at a value

of M and then trying again. This must be continued until left handed side (LHS) value of

equation(3) and right handed side (RHS) value of equation(3) should yield same result.

RESULT

The values from Graphical method and Analytic method should be equal in two

decimal places.


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PART - II

AIRCRAFT PERFORMANCE

EXPERIMENT 7

STATIC PERFORMANCE: THRUST REQUIRED CURVE

AIM

To conduct static performance analysis using thrust required curve. Hence calculate

the following:

1) Minimum thrust required ( r T min),

2) Thrust available ( aT ),

3) Maximum velocity ( ma xV ).

REQUIREMENTS Wing area Aspect ratio Drag polar Span efficiency factor.

THEORY

The performance of an airplane for an uncelebrated flight conditions is called static

performance.

Thrust required ( r T ) is given by

D LW

C C W

T

D L

r

Thrust required curve is a plot of the variation of r T with respect to velocity V

PROCEDURE

To calculate a point on the thrust require curve

1) Choose a value of V

2) For this V , calculate the lift coefficient LC from equation.

3) Calculate DC from the known drag polar for the airplane.


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AeC

C C Ld D

2

0

4) Calculate the ratio d L C C .

5) Calculate thrust required from equation (1)6) The value of r T obtained from step five is that thrust required to fly at the specific velocity

taken in step1. Fig1 is the locus of all such points taken for all velocities in the flight range of

the airplane.

7) The thrust required will be minimum when zero lift drag due to lift.

di L

d C AeC

C

2

0

Maximum Thrust available: It is the maximum thrust provided by an engine-propeller/jet

Maximum Velocity: The velocity of airplane obtained when thrust available is maximum


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EXPERIMENT 8

STATIC PERFORMANCE: POWER REQUIRED CURVE

AIMTo conduct static performance analysis using power required curve. Hence calculate the

following:

1) Power available ( r P ),

2) Maximum velocity ( ma xV ).

REQUIREMENTS Wing area

Aspect ratio Drag polar Span efficiency factor.

THEORY

The performance of an airplane for a uncelebrated flight conditions is called static

performance.

Thrust required ( r T ) is given by

D LW

C C W

T D L

r

Power required is given by

V T P r r

Power required curve is a plot of the variation of r P with respect to velocity V


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POWER AVAILABLE:

a) PROPELLER:

Shaft Break Power (P): The power delivered to the propeller by the crank shaft is

defined as the shaft break power. Not all P is available to the drive the airplane, some of it

dissipated by inefficiencies of the propeller itself. The power available to propel the aircraft

Pa is given by

P P a

b) Jet:

The power available from the jet engine is obtained from

V T P aa

Maximum Velocity : The velocity of airplane obtained when power available is maximum

PROCEDURE

To calculate a point on the thrust require curve

1. Choose a value of V

2. For this V , calculate the lift coefficient LC from equation.

3. Calculate DC from the known drag polar for the airplane.

AeC

C C Ld D

2

0

4. Calculate the ratio d L C C .

5. Calculate thrust required from equation(1)

6. The value of r T obtained from step five is that thrust required to fly at the specific

velocity taken in step1.

7. Calculate the power required using the equation V T P r r

8. The power required curve is defined as a plot of r P versus V as shown in Fig1

Documents

FVD Lab Manual