hr uas report

8/3/2019 Hr Uas Report

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HUMANATARIAN RELIEF-UNMANNED AERIAL

SYSTEM (HR-UAS)

Aerospace Vehicle Design

Project Report

Submitted by

Amit Nayak

SC08B062

Yogesh Khedar

SC08B104

In

Department

Of

Aerospace Engineering

Indian Institute of Space Science and Technology

Thiruvananthpuram


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LIST OF FIGURES

Figure 1 Mission Profile

Figure 2 Technoavia SM-92 Finist

Figure 3 Gippsland GA-8

Figure 4 De Havilland DHC-2

Figure 5 Cessna 421

Figure 6 Design Methodology Followed

Figure 7 Empty weight Fraction Estimates

Figure 8 Engine types used in relation to the Speed-Altitude Envelope of Aircrafts

Figure 9 Payload vs GTOW trade off

Figure 10 Solution of Constraint Analysis

Figure 11 Effect of on Figure 12 Thckness Ratio Historical Trend

Figure 13 Plain Flap

Figure 14 Airfoil Profile of the NACA 23018

Figure 15 vs for NACA 23018Figure 16 Hoerner Wing Tip

Figure 17 Tail configurations for spin recovery

Figure 18 Ballistic Recovery System in use

Figure 19 A functional BRS system

Figure 20 Fuselage Sizing parameters

Figure 21 Wheel load geometry


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LIST OF TABLES

Table 1 Requirement Captured

Table 2 Comparable aircraft Characteristics

Table 3 Mission weight fraction estimates

Table 4 Initial Sizing Results

Table 5 Pusher Configuration

Table 6 Tractor Configuration

Table 6 Canard

Table 8 Wing Placement options

Table 9 final wing geometry after refined constraint analysis and refined sizing:

Table 10 Vertical tail

Table 11 Horizontal tail

Table 12 Single vs Double

Table 13 Engine Characteristics:

Table 14 Number of blades selection for Propeller

Table 15 Empty weight fraction parameters

Table 16 Refined Sizing Results

Table 17 Parameters for Fuselage Length sizing

Table 18 Parameters for fuselage sizing

Table 19 Tail volume coefficients for general aviation plane

Table 20 Summary of refined sizing

Table 21 Weight Breakup

Table 22 Location of Center of Gravity from nose


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Table 23 Location of neutral point from nose

Table 24 Location of leading edge from nose

Table 25 Loads on landing gear

Table 26 Tire data


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Humanitarian Response Unmanned Aircraft System (HR-UAS)

Introduction:

Unmanned aerial vehicles have been operating in various roles for a number of decades.They have been most extensively used for providing intelligence, surveillance, and

reconnaissance and now as strike platforms. There is no doubt that this platform has enormous

potential in serving many functions and one in particular has gained a lot of interest, especially in

the wake of recent natural disasters like the earthquakes that struck in japan and Haiti, is the

cargo resupply for humanitarian assistance missions.

There is a need for an affordable, humanitarian response aircraft system that can provide

aid to the populations of both developed and under-developed nations worldwide when natural

disasters occur.

These natural disasters strike with no warning and hence there exists a need for a high

level of readiness in rescue operation also, logistical operations are often strained by the

magnitude and nature of these disasters. Precision, unmanned resupply solutions could help

alleviate some of these problems. Use of such a precision unmanned system would allow for

delivery of critical supplies to remote, discontinuous areas where terrain and environment have

limited the access of current assets and also reduce potential for causalities while serving in

hostile or politically sensitive areas.

This report presents the development of a Humanitarian Response Unmanned Aircraft System(HR-UAS) which can achieve all objectives as and when demanded by the above mentioned

situations. The HR-UAS is planned to be used in India and hence all requirements have beenprimarily influenced by the unique operating environments existing in India.


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Requirement formulation:

This section will report the decisions taken concerning the possible mission requirements

of the planned unmanned vehicle (HR-UAS).Since this aircraft is to be operated in India, a

recent eventThe Sikkim Earthquake has been studied so that important lessons learnt from this

experience could be implemented in the design of this aircraft.

The Sikkim Disaster:

The earthquake struck Gangtok with a magnitude of 6.9.The relief operations started

immediately. Relief operations to South and West Sikkim were delayed due to landslides. Hence

air assets were immediately pushed to carry out these operations. Bagdogra airport was made thecentre of relief operations. All relief materials would reach this airport and would be transported

via various means to the targeted area. The affected areas included ones which were very remote.

The relief operations had to rely on air based assets for airdropping/delivering relief packages to

the needy.

Cheetah helicopters were used for recce missions. C-130J Hercules aircrafts were used to bring in relief materials and NDRF personnel to

Bagdogra airport.

IL-76 aircrafts were also used for similar operations. Mi-17 helicopters were used to carry medical personnel to tend to the wounded in the

affected areas.

ALH Dhruv also performed a similar role. Remaining air assets; mostly helicopters were employed for dropping/delivering relief

materials to the needy.

Lessons learnt:

-Helicopters were used to deliver the relief materials and this could be done only after recce

missions by Cheetah helicopters. HR-UAS would have been more effective in such a situation as

it is capable of doing both recce and airdropping of materials simultaneously.

-Helicopters could have been put to better use eg: evacuating, bringing in personnel and air-

ambulance roles.

-The helicopters have a range of 450-500 km which in these situations could be a serious bottleneck.HR-UAS on the other hand being a fixed-wing aircraft promises both high range and

responsiveness.

-Also risky missions like these put the crew involved in danger, thereby increasing the chancesof causalities.HR-UAS being an unmanned system would eliminate such risks.


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-No fixed wing aircraft were used for air dropping purposes which means that the capability to

conduct such operations is non-existent in India.


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REQUIREMENTS CAPTURED:

Having considered the airport/airstrip density in India, we can safely conclude that an

airstrip exists within 600km of another. Hence the maximum ferry range of HR-UAS should be

1200km.The aircraft should have a service ceiling that can enable it to operate from any airstrip

in India. Leh airport is located at a height of 4000 MSL and hence HR-UAS should have aceiling of around 5000MSL.The aircraft should have a very small unit price cost, life cycle cost

and be reliable in order to make it an attractive option reserved for such operations. Small utility

& general aviation aircrafts offer such qualities. Also instead of having a very large aircraft itsbetter to have a small aircraft which can offer more precision and better reach. The aircraft

should be designed in the lines of the best in the general aviation and utility aircraft category.

Payload has been set as 850kg to match the payload carrying capacity of majority of the aircraftsin the above mentioned categories. The smallest airstrip in India is at Along in Arunachal

Pradesh; measuring 400m.The minimum landing length for HR-UAS should therefore be

400m.The cruise speed for HR-UAS has been decided to be 250kmph based on the values of the

same for all utility and general aviation aircraft which the HR-UAS plans to match in terms of

flight performance.

The aircraft will airdrop the relief material to the target areas to avoid landing in difficultterrains and in hostile places. Any human involvement imposes many additional constraints on

the vehicle and the mission itself. Hence to limit human involvement in the vehicles operationsit is planned to have an autonomous systems capable of waypoint navigation en route and human

in the loop takeoff and landing. A support team will be deployed in the airstrip to conduct theoperations of the HR-UAS .To minimize the effort and cost required to deploy such an unit the

ground station should be compact and light. Also the aircraft should be light and compact so that

it can be carried by a heavy transporter in case such a need arises. The unit cost, complexity,operation readiness time and maintenance cost should be as low as possible. These are the

driving factors of the design process.

Table 1 Requirement Captured

Requirement

Payload 850kgs

Ceiling 5000MSL

Range 1200km

Cruise Speed 250kmph

Take off Distance


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MISSION PROFILE:

Based on the captured requirement a mission profile is formulated. The mission profile is

illustrated in the given figure and displays entire mission from takeoff to landing. The aircraft

will take-off from the operational airport climb to the cruise altitude. It will cruise to the disaster

hit region, locate the dropping point and descent to a drop altitude. It will then drop the material,climb back to cruise altitude and return to the operational base.

Figure 1 Mission Profile

1-2 Taxi and Take-off

2-3 Climb to cruise altitude

3-4 Cruise to destination

4-5 Decent to air drop altitude

5-6 Air Drop

6-7 Climb to cruise altitude

7-8 Cruise back to Forward base

8-9 Descent for landing


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9-10 Loiter if required at loiter altitude

10-11 Descent to land

11-12 Land and taxi

COMPARABLE AIRCRAFTS

UAVs of this weight category and function does not exist at present. But there are lots of

General Aviation planes exist of this class which can be compared for their characteristics. Of

various aircrafts compared following 4 were close to our requirements:

Figure 2 Technoavia SM-92 Finist

Figure 3 Gippsland GA-8


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Figure 4 De Havilland DHC-2

Figure 5 Cessna 421


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Table 2 Comparable aircraft Characteristics

Characteristic SM-92 GA-8 Cessna 421 DHC-2

Propulsion Piston-Tractor Piston-Tractor Piston-Twin Piston-Tractor

Payload 600 950 800 950

GTOW 2350 1800 3400 2300Range 1400 1400 2800 730

Cruise Speed 200 220 440 230

Ceiling 3000 6000 9000 4600

Mission Type STOL/UTILITY UTILITY PERSONAL STOL/UTILITY


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INITIAL SIZING

This section breifly explains the method used for initial sizing and various desicions and

trade offs done during the process to arrive at a GTOW estimate.The figure below is flow chart

of the series of analysis that one must conduct in order to arrive at a GTOW estimate.

Figure 6 Design Methodology Followed


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To arrive at the initial estimate of GTOW (Gross take off weight) the following equation was

used,

Which can be written as,

where

Starting from a guess value of and iterating we arrive at a initial guess of GrossTakeoff Weight (GTOW).

Empty Weight Fraction:

Figure 7 Empty weight Fraction Estimates


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Empty fraction () is taken from above table for a single or double engine general

aviation aircraft as this category best resembles our aircraft.

Fuel Fraction Estimates:

The following table gives weight fractions for different parts of the mission profileswhich are quite accurate for the purpose of initial sizing.

Table 3 Mission weight fraction estimates

Mission Segment Warm-up and Takeoff 0.970

Climb 0.985

Landing 0.995

Cruise segment mission weight fractions are found by using Breguet range equation:

Where,

Loiter weight fraction is found from the endurance equation:

Where,

Both of the above equation requires for calculation of weight fraction and since no

configuration was available at this stage we decided to choose this ratio based on historical trend.

Aspect ratio was selected to be 6 purely on the basis of historical trend of similar systems. Also


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for an AR that is between 3 and 7 historical data predicts to be twice this value so wehave chosen ( . From this value of ( the ( for different sections of themission profile can be found out.

PROPULSION SYSTEM SELECTION:

Figure 8 Engine types used in relation to the Speed-Altitude Envelope of Aircrafts

Based on the flight envelope outlined in the mission requirement we selected Piston prop as our

propulsion system. Decision on number of engine will be taken at a later stage.


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Figure 9 Payload vs GTOW trade off

Following are the results obtained from initial sizing for two configurations namely,

single engine and double engine aircraft. A sensitivity analysis was done based on the results

below and is shown above. GTOW(1) refers to the single engine aircraft and GTOW(2) refers to

the double engine configuration aircraft.

Table 4 Initial Sizing Results

PAYLOAD GTOW (SINGLE -

ENGINE)

GTOW(DOUBLE -

ENGINE)

400 1100 1470

600 1500 2050850 2000 2720

0

500

1000

1500

2000

2500

3000

0 200 400 600 800 1000

G

T

O

W(

k

g

s)

Payload(kgs)

Payload vs GTOW

GTOW(1)

GTOW(2)


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CONSTRAINT ANALYSIS:

The following requirements are affected by W/S.

Landing Ceiling Stall velocity

Landing:

Landing distance is largely determined by wing loading. Wing loading affects the

approach speed. The approach speed determines the touchdown speed, which in turn determines

the kinetic energy which must be dissipated to bring the aircraft to halt. Hence smaller the W/S

we have the smaller the landing distance will be.

Equation given below provides a good approximation of the landing distance, which is

used to estimate the maximum landing wing loading.

Where,

Ceiling:

Wing loading (W/S) from ceiling constraint is found from the following equation.

Where,


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Ceiling puts an upper limit on the wing loading.

The following requirements are dependent on both W/S and T/W.

Take-off Distance Turn rates Range and endurance Climb performance

Requirements which are dependent solely on T/W are:

Missed approach gradient Climb gradient

The above requirements for T/W are usually set by FAR regulations and since HR-UAS is anunmanned system we wont have any such constraints.

Range:

To maximize range during cruise, the wing loading should be selected to provide a high at the cruise conditions. A propeller aircraft, which loses thrust efficiency as speed goes up,gets the maximum range when flying at the speed for best . Therefore, to maximize range apropeeler aircraft should fly such that

During cruise, the lift equals the weight, so the lift coefficient equals the wing loading

divided by the dynamic pressure. Substitution into the above equation allows solution for the

required wing loading to maximize for a given flight condition.

Loiter:

In line with other civil aircrafts we have fixed the loiter requirement during the mission to

be 20min before landing (Even unmanned systems need to wait for Air Traffic Controllers(ATC)

affirmation before landing). Following relation gives wing loading for optimum loiter

performance,


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Takeoff:

Both wing loading and thrust-to-weight ratio(equivalent) contribute to the take-off

distance. Following relation is used for finding the relation between wing loading and thrust-to-

weight ratio (equivalent T/W in this case):

Rate of climb:

Rate of climb is the vertical velocity. Climb gradient, G, is the ratio between the

vertical distance travelled and the horizontal distance travelled. After some manipulation with

various parameters following relation is obtained between Wing loading and (equivalent):

* + * +

Where,

Result of constraint analysis is plotted in the graph below. From the graph a point (T/W,W/S) is

chosen so that we have minimum T/W. Its cheaper to invest in stronger wings then to go for a

power plant that has a higher power rating and consumes a lot of fuel.

Wing loading

Power-to-weight ratio


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Figure 10 Solution of Constraint Analysis


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CONFIGURATION LAYOUT

Propulsion system configuration: For a piston prop engine; Pusher and tractor configurations

are the most widely used. The advantages and disadvantages of each configuration have been

discussed briefly and the most relevant configuration with respect to HR-UAS has been chosen.

Pusher vs Tractor:

Table 5 Pusher Configuration

Pros Cons

This makes the propeller slightly more

effective because it receives relatively slowerair and the prop-wash is unimpeded.

Additional engineering efforts are usually

required to prevent dragging a tail-mounted(pusher) prop during takeoff and landing.

Allows special equipment (radar, AUVcameras) to be efficiently installed in the

fuselage nose.

A pusher design with an empennage behind thepropeller is structurally more complex than a

similar tractor type.A smaller vertical tail During pitch rotation at takeoff, to avoid

ground contact propeller diameter may have to

be reduced (with a loss of efficiency) and/orlanding gear made longer and heavier.

Less form drag This reduces propeller efficiency and causesvibration inducing structural propeller fatigue

and noise.

Wing profile drag may be reduced due to the

absence of prop-wash over any section of the

wing.

In pusher configuration, the propeller does not

contribute airflow over the engine or radiator.

Allows shorter fuselage. CG in the aft region is not good forlongitudinal stability.

Longer runway lengths required.

Table 6 Tractor Configuration

Pros Cons

The heavy engine is at the front, which helps to

move CG forward and therefore allows asmaller tail for stability considerations

The Propeller slipstream disturbs the quality of

airflow over fuselage and wing root

Propeller is working in undisturbed free stream The increased flow velocity over fuselage andturbulence increases the skin friction drag of

the plane

There is more effective flow for engine cooling

Tractors are naturally stable. Thrust comingfrom the most forward part of the airframe has

the inherent tendency to stay in line of the

direction of travel. There is no substitute for


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the stability gained in having the propeller up

front.

After the constraint analysis the power required for HR-UAS was found to be 400 HP.

The large power requirement will force us to use a large piston prop engine. Large piston propshave cooling issues and these problems further aggravate in a pusher prop configuration which

offers very low air cooling as free stream air is not available. The pusher configurations with

large engines require auxiliary fans to cool the engine. This would increase complexity of the

system and hence the maintenance cost and would adversely affect the reliability of the system.

From the point of view of one of our primary requirements that is air dropping of relief

material, we have two choices 1) to drop the packages from the belly of the aircraft 2) to release

the packages from the aft end of the plane. Both these choices offer a few advantages and

disadvantages however the 1st option where the relief material is to be dropped from the belly

(much like a heavy bomber aircraft) imposes certain problems relating to the structure of thefuselage, landing gear and camera mounting location. Such a configuration would also be

favored by a pusher prop configuration however it will result in thickening of the mid fuselage

structure (increasing the weight), Camera mounting which should be done so as to enable a

panoramic view of the terrain will have to be done at a location other than the belly which would

have given the best results. Having considered the above points it was decided that dropping the

relief material from the aft end was a better option and hence for the same reason the pusher prop

configuration was dropped which would have not favored this mode of air dropping.

If we consider stability to be one of the core concerns then also tractor favors well then

pusher which tends to make the system slightly unstable in longitudinal stability.

These arguments led us to select tractor as our basic configuration over pusher.

Wing and Tail configuration:

Second very important decision in configuration layout was selection wing and tail

configuration. In this case also of the various options were available of which two were

shortlisted for serious investigations. These are:

1. Canard configuration (Tandem wing)2.

Conventional

3. T-TailTable 6 a) Canard

Advantages Disadvantages

Good stalling characterisitcs

Helps if pusher is selected. CLmax problems with flaps or margin on the


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entire wing: Flaps produce a larger pitching

moment about the c.g. on a canard aircraft.

Synergistic use of winglets for directional

stability.

Induced drag / CLmax incompatibility

Provides positive lift Directional stability:May require large VT

Flaps may not be required The wing additional load distribution isdistorted by the canard wake.

Power effects on canard - deep stall: If canard

stalls before wing.

Control authority is larger for unstable

canard aircraft at high CL than for unstable

aft-tail designs.

It is easy to make a very bad canard design.

This may require larger wing area then

required.

Takeoff and landing distances and speeds are

often higher than for similar conventionalaircraft.

The wing root operates in the downwash fromthe canard surface, which reduces its

efficiency,

Table 6 b) Conventional vs T Tail

Conventional Tail T Tail

More predictable design characteristics Maintenance issues

Better pitch control especially at low speeds. Very dangerous deep stall may be encountered

Vertical stabilizers must be made strongeradding weight.

After deliberate discussions on the above three options we found that it was not worth

taking the risk to use either canard configuration or T-tail configuration for our design as their

cons overweighs their pros. Also the benefits which these configurations seem to be providing

over conventional layout were not convincingly significant for the HR-UAS objectives. We

selected the conventional tail configuration for HR-UAS.


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AERODYNAMICS:

The following aspects play a pivotal role in making the aircraft aerodynamically efficient.

Wing Placement Airfoil selection Wing configuration Wing parameters

WING PLACEMENT:

Three options are available for the vertical placement of the wing with respect to the fuselage.

Table 8 Wing Placement options

High Mid Low

Allows placing fuselagecloser to ground, thus

allowing loading and

unloading without specialground handling

equipment.(This is veryimportant for HR-UAS)

Advantages of groundclearance as in the high wing

configuration.

To provide adequate groundclearance, the fuselage has to

be at a higher level

as compared to the high wingconfiguration.

Fuselage generally houses

the landing gear in

special pods leading to higherweight and drag.

Lower drag due to reduced

wing -body interaction.

Landing gear can be located in

the wing thereby

avoiding pods on the fuselageand hence lower form

Drag.

For short take off and landing

(STOL) airplaneswith high wing configuration

have following specific

advantages. (a) Large wingflaps can be used (b)

Engines are away from the

ground and hence

ingestion of debris rising

from unpreparedrunways is avoided (c)

Prevents floating of wingdue to ground effect which

may occur for low wing

configuration.

The mid wing placement of

wings is more of a trade off incertain situations. It enables to

enjoy the benefits of both the

low wing and high wingplacement of the wings.

Takes better advantage of

ground effect.

Dihedral not needed in this

case due to inherent roll

Neutral stability. A low-wing configuration has

unstable contribution


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stability offered by high

wing.

to the directional stability.

Hence a larger vertical

tail area is needed.

For low speed airplanes,

weight saving can beeffective by strut braced wing

assembly.

Structure weight is least

Wing mounting is complex as

the wing often has to go

through the fuselage. This maybe an obstruction for the

payload compartment.

Max Structure weight and

complexity

Easier to make cantilever

(strutless) wing without

sacrificing room for payload.Structure weight is least

The above comparisons have fallen in favor of the high wing configuration when design

requirements of HR-UAS are kept in mind. HR-UAS will thus have a high wing with no

dihedral. Sweeping the wing was ruled out as it offered no significant benefits for the flight

envelope that HR-UAS is being designed for. Tapering of the wings helps in reducing induced

drag and also results in weight reduction. But when costs are taken into consideration, theincrease in manufacturing costs outweigh the weight losses (Unit Cost of HR-UAS will be

instrumental in making it an attractive product for relief operations).For the above reason a

rectangular wing design will be adopted.

AIRFOIL SELECTION:

The airfoil affects the cruise speed, takeoff and landing distances, stall speed, handling

qualities (especially near stall), and overall aerodynamic efficiency during all phases of flight.

There are primarily 3 types of NACA airfoils available. While rarely used for wing

design today, the uncambered 4-digit airfoils are still commonly used for tail surfaces of

subsonic aircraft. While the 5-digit airfoils offer greater maximum lift by shifting maximum

camber forward, the 6-digit offer increased laminar flow and hence decreased drag. Hence the

above are the options for the choice of airfoils for HR-UAS.


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Figure 11 Effect of t/c on Clmax

Figure 12 Thckness Ratio Historical Trend

The airfoil chosen for the proposed aircraft wing is NACA 23018. The decision is based

on the fact that out of the four aircrafts which closely resembled HR-UASs design, two of themused this airfoil. This airfoil is a thick laminar airfoil. Thicker section allows it to have higher

structural stiffness to take higher wing root bending moment loads. Laminar flow allows it to

have a lower drag coefficient under various flight conditions. Additionally a thicker airfoil

allows having a higher volume for storage of fuel in the wing (Increasing more payload volume).

NACA 23018 was chosen as the airfoil for the proposed aircraft after having considered the

above benefits.


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To aid in the takeoff and landing performance, we plan to have a wing with trailing edge

flaps. For simplicity (and hence, production cost savings and maintenance costs), we choose a

simple plain flap and no slats.

Figure 13 Plain Flap

Such a flap when deflected by 45 deg will yield an increase in the airfoil maximum lift

coefficient by 0.9; which is substantial. Hence, for NACA 23018 airfoil employing theabove flaps, the maximum lift coefficient will be 2.5. with 45 deg flap deflection=1.6+0.9=2.5

To account for the three-dimensional effects of a finite wing, the wing lift coefficient is given by

Figure 14 Airfoil Profile of the NACA 23018


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Figure 15 Cl vs Alpha

TWIST:

Any attempts to optimize the lift distribution by twisting the wing will be valid only at

one lift coefficient. At other lift coefficients, the twisted wing will not get the whole benefit of

twist optimization (D.P.Raymer,Aircraft Design: A conceptual Approach Fourth

Edition,).Therefore no twist was given.

WINGLETS:

The most widely used low-drag wing tip is the Hoerner wing tip. This is a sharp-edged

wing tip with the upper surface continuing the upper surface of the wing. The lower surface is

undercut and canted approx. 30 degree to the horizontal. The lower surface may also be undercambered(D.P.Raymer,Aircraft Design: A conceptual Approach Fourth Edition,).The Hoerner

wing tips increase the distance between the vortices developing from the wings. This reduces the

wing tip drag dramatically. Hoerner wing tips provide the largest effective span for a given

geometric span or a given wing weight. The effect of these winglets is best illustrated in the

following figure.


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TAIL GEOMETRY AND ARRANGEMENT

Tail Arrangements:

H Tail-To position the VT in undisturbed air during high AOA conditions or to position

the rudders in the prop-wash on a multiengine aircraft to enhance engine out control. The H Tailis heavier than the conventional tails but its endplate effect allows having a smaller HT.

It also allows having smaller height of the tail for ease in maintenance and transportation.

Spin recovery:

VT plays an important role in spin recovery. Enough rudder control is required at high

AOA to stop the rotation of airplane about vertical axis and reduce the excessive side-slip angle.

At high AOA the HT is stalled, producing a turbulent wake extending upward approx.. 45 degangle. As a rule of thumb 1/3 of rudder should be out of wake of tail.

Dorsal and Ventral Fins: The dorsal fin improves tail effectiveness at high angle of side

slip by creating a vortex that attaches to the vertical tail. This helps in preventing high angle of

sideslip even in spin, and augments in rudder control in the spin. The ventral tail also tends to

prevent high sideslip, and has the extra advantage of being where it cannot be blanketed by the

wing wake.

Figure 17 Tail configuration for spin recovery


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The exact plan form of the tail surfaces is actually not very critical in the early stages of

the design process. The tail geometries are revised during later analytical and wind tunnel

studies. For conceptual design it is usually acceptable to draw tail surfaces that look right based

on prior experience and similar design.

For low speed aircraft, there is little reason for VT sweep beyond 20 degrees other thanaesthetics.

In addition to the various studies and comparisons conducted with other aircrafts of

similar weight categories, V-tail configuration has been abandoned. The V-tail empennage

configuration was also abandoned for similar reasons. V-tails are very complex control surfaces

and involve coupling yaw and pitch control and could make manual control (unassisted by

computers) of the UAV difficult. Although the V-tail may reduce tail weight by reducing the

need for multiple surfaces, simplicity is the key objective with the aircraft. The V-tail also has

trouble maintaining longitudinal and directional stability and is susceptible to Dutch roll. Withthat being said the V-tail has been determined to not be useable in this design.

The H-tail configuration was driven by the need to increase stability in terms of spin

resistance. By having two vertical stabilizers, the area of the individual vertical stabilizer is

reduced. This decreases the side profile of the UAV so that in case of a sudden cross wind, only

one of the vertical stabilizers will be affected allowing the other to apply a correcting maneuvers

that will help stop or reduce the spin of the UAV. Also the horizontal stabilizer will be directly in

the flow from the engine making the elevator more effective during flight. However, the use of

three surfaces will increase the weight of the tail section which will increase the overall weightof the UAV.

TAIL SIZING:

Tail airfoil selection:


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Unlike wing, whose function is to generate lift strong enough to sustain the airplane in

the air, the aerodynamic forces generated on the tail are relatively small; they need only be large

enough to maintain the stability and control. Also the aerodynamic forces on the tail readily

change directions depending upon various flight conditions like yawing right or left and/or

pitching up and down. Hence it makes no sense to use cambered airfoil for tail sections; rather

the horizontal and vertical tail on almost all airplanes use symmetrical airfoil section. A popular

choice is the NACA 0012 airfoil. We have decided to use the same airfoil for both horizontal and

vertical tail to retain commonality and hence cost reduction.

Control surface sizing:

Horizontal Tail: Wings of low AR although are aerodynamically less efficient, stall at a

higher value of angle of attack than wings with higher AR. Hence, if the horizontal tail has a

lower AR than the wing, when the wing stalls, the HT still has some control authority. To

achieve this advantage in good controllability in stall, we select a AR for tail to be lower than the

wing. From historical point of view the AR of HT for GA planes has been between 3-5. So wechoose the average value of 4 to be consistent with the historical dat. To be consistent with the

wing we chose taper ratio of 1. The span of the wing is found from,

The tail root chord is given by,

And tail tip chord is given by,

The spanwise location of the mean aerodynamic chord for the HT is,

VT- Typical aspect ratios for vertical tails range from 1.3 to 2.0, where the aspect

ratio is based on the root-to-tip height

(as span from tip tip does not have any relevance

here).

We choose an aspect ratio of 1.5 again on historical basis. The above equation gives the value of . The taper ratio given is 0.5. The root chord is given by,


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The tip chord then can be given by,

Again the vertical location of the mean aerodynamic chord of the vertical tail, referencedto the root chord, is

TAIL SIZING RESULT:

Table 10 Vertical tail

Aspect Ratio() 1.5Taper Ratio() 0.5Root chord length() 1.24mTip chord length() 0.62mVT reference area() 2.6 m2Vertical location of mean aerodynamic

chord( ) 0.617 mTail height 1.39 m

Table 11 Horizontal tail

Aspect Ratio 4Taper Ratio () 1Root Chord length () 1.376mTip chord length () 1.376mHT reference area () 7.584 m2Spanwise location of mean aerodynamic

chord

)

1.3769 m

Tail span 5.5078 m


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PROPULSION

Table 12 Single vs Double

Single engine Twin engine

Low maintenance costs Maintenance costs almost doublesLow Unit cost High unit cost(>2xcost of single engine plane)

Life cycle cost is lower owing to the above two

factors

Taking into account above factors the life cycle

cost for this option is higher

Higher GTOW then single engine(more than

1.5 times)

Historical data by NTSB (National

transportation security board) has proved that

single engine accident rates fare comparably

with twin engine planes.

Reliability increases but at the same

P/W for each engine should be enough to

enable the aircraft to land safely > 50% of total

power.

Will have lesser drag Will have more drag

Cost can be minimized by limiting the size of the plane so that a power plant

configuration can be employed which is reliable, requires minimum maintenance and offers a

better price (this will depend heavily on its availability or demand). Single engine option gives us

most of these advantages over their double engine counterpart.

Ballistic Recovery System

Reliability is one of the main reasons why double engine is preferred over single engine

configurations by customers even if they may be cost more. To enhance reliability whileretaining the advantages of single engine option we propose to use a Ballistic Recovery system

(parachute recovery). Though this system is new and remains untested for the weight class of

planes that is represented by HR-UAS, we believe such a system could radically improve the

safety of this asset without much increase in price. This system recovers or brings down the

failed aircraft to ground safely with the help of a parachute. This system has a very minimal

weight penalty (Weight increase


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Figure 18 Ballistic Recovery System in use

The above figure shows the sequence of events involving an aerobatic aircraft breaking

under high loads. This could have easily turned into a tragic incident. The pilot had the BRS

system to thank for as he walked out of the incident with only a few bruises.

Figure 19 A functional BRS system

The above figure shows a BRS system for a light aircraft. The system is quite sleek and

light.

Weight of this system for an aircraft of GTOW ~ 500 kgs is around 10 kgs. Assuming a

linear variation of weight of the system and GTOW, we get a weight of 45 kgs for our system.

No readymade BRS is available for this kind of GTOW but many companies are willing to offer

customized solutions for various weight categories, besides we have a national laboratory (of

DRDO) working in this area, so to develop a system for our vehicle wont be difficult.


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Engine Selected: Lycoming IO-720A

Table 13 Engine Characteristics:

Name Power RPM Height Width Length Dry Weight

IO-720A 400HP 2650 572.2mm 869.95 mm 1173.48 mm 272.4 kgs

Propeller sizing:

The actual details of the propeller design, such as the blade shape and twist, are not

required to lay out a propeller-engine aircraft. But diameter of the propeller, the dimensions of

the engine, and the required inlets and exhausts must be determined. Since limitation on

propeller diameter is the tip speed, we computed the tip speeds for 2 and 3 bladed propeller.

Comparing the two we selected 3 bladed propeller since its tip velocity is much lower than sonic

speed. Diameter is calculated from following relation,

The tip of a propeller follows a helical path through the air. Tip speed is the vector sum of a

rotational speed and the aircrafts forward speed. Therefore,

Where

Net Helical speed is given by,

Table 14 Number of blades selection for Propeller

()

()

2Blade

0.56 2.327 321.36 70 328.90 335

3Blade

0.52 2.180 302.458 70 310.458 335


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41

MOUNTING

Cooling:

Cooling is a major concern and especially so when using a powerful piston prop

(>300hp). Up to 10% of the engines power can be lost by the drag associated with taking incooling air, passing it over the engine, and ejecting it.

To minimize this cooling drag, the cooling-air mass flow should be kept as small as

possible while maintaining engine temperature to optimal levels. Optimization studies indicate

that the best intake slows the air 30% of the aircraft flight speed (climb speed in the worst case).

The above studies result in the following equation for piston engine cooling sizing:

Where Power is in kW and climb velocity in .


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REFINED SIZING

Rubber engine sizing is used during early stages of aircraft development program where

either a new engine is warranted or where numerous options for engine in various power

increment categories are available. The latter was the case for our design. There are numerous

options available for piston-prop engine of power less than 500hp. Rough sizing resulted in

engine power of around 380hp. So it was assumed that the power requirement cannot go higher

than 500hp and hence rubber engine sizing was done.

Also rubber engine sizing approach allows us to size the aircraft to meet both the

performance and range goals i.e. while the takeoff weight changes so does the power

requirement. The rubber engine is then scaled accordingly.

Refined sizing equation:

As in the initial sizing case, an initial guess of the take-off weight is used (result of initial sizing

is used in this case) to determine a refined take-off weight, and the solution is iterated until the

two are approximately equal to within a small error margin. Difference in refined methods and

the initial sizing method is that in the former the empty-weight fraction is calculated and used .

Empty Weight Fraction:

The empty weight fraction is estimated using improved statistical equations which reflect the

weight impact of the major design variables in a more effective way. These variables are aspectratio (A), power-to-weight ratio (P/W), wing loading (W/S) and maximum speed.

The equation used for a propeller aircraft is:

Values for various constants taken for computation of empty weight fraction are given below.

Table 15 Empty weight fraction parameters

Fps units GA-

SINGLE

-0.25 1.18 -0.20 0.08 0.05 -0.05 0.27


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REFINED SIZING RESULTS:

The following is the tabulation of the iterations done in order to arrive at the refined

GTOW(W0).

Table 16 Refined Sizing Results

GUESS W0 We/W0 W0,CALCULATED in

pounds

4400 0.5138 4764.6

4764.6 0.5018 4894.5

4894.5 0.4977 4939.9

4939.9 0.4964 4955.7

4955.7 0.4959 4961.2

4961.2 0.4957 4960(2250 kgs)

Fuselage sizing:

Initial estimate of fuselage length can be found from following relation which is a

statistical equation based on historical trends. These are based solely upon GTOW. The

corresponding constants used are presented in the table below.

Figure 20 Fuselage Sizing parameters


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Table 17 Parameters for Fuselage Length sizing

SI Units General Aviation-Single 1.6 0.23

Table 18 Parameters for fuselage sizing

Aircraft Type GA-Single Engine 5-8(6) 3-4(4) 3-9(9)

Values in the bracket are selected values for computation of various dimensions.

Tail volume coefficient:

For initial layout, a historical approach is used for the estimation of tail size. The

effectiveness of a tail in generating a moment about the center of gravity is proportional to the

force(i.e. lift) produced by the tail and its moment arm.

Tail volume coefficients are given by,

Where the moment arm is approximated as the distance from the tail quarter chord to thewing quarter chord.

Values of tail volume coefficients were taken from historical data and are given in the table

below,

Table 19 Tail volume coefficients for general aviation plane

Horizontal () Vertical ()GA-Single 0.70 0.04


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Control surface sizing

Based on information available in textbook (Raymer), aileron has been taken to extend

from 50% span to 90% span. Flaps occupy the part inside of ailerons.

Table 20 Summary of refined sizing

GTOW 2250KGS

Fuselage Length 9.443 mVertical Tail area 2.600m2Horizontal Tail Area 7.584 m2


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WEIGHTS AND BALANCE ANALYSIS

To find weight breakdown of various components of the vehicle a Statistical Group

Weights Method (SGWM) was used, which gives a refined estimate of the group weights basedon sophisticated regression analysis. It should be understood that there are no right answers in

weight estimation until the first aircraft flies. However these equations provide a reasonable

estimate of group weights.

The result of this analysis is tabulated below:

Table 21 Weight Breakup

Group Weight(in lbs) Weight (in kgs)

Wing 575 260

Horizontal Tail 87 40Vertical Tail 57 25

Fuselage 725 330

Installed Engine 940 425

Fuel System 80 35

Flight Control System 100 45

Avionics 43 20

Electrical 145 65

Landing Gear 285 130

Payload 1870 850

Total 4907 2225

Using values obtained from weight analysis center of gravity of the aircraft was found

using an elementary approach given by Anderson, Aircraft Performance and Analysis. The

results obtained are tabulated,

Table 22 Location of Center of Gravity from nose

Loaded Unloaded

CG from Nose(m) 3.77 2.45

These values are used to locate the wing on the fuselage. The following relation between

the location of aerodynamic center of the wing body combination and the location of the neutral

point is presented in Introduction to Flight by Anderson,


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Where,

Now for positive static stability static margin should be positive. For conventional generalaviation airplanes, the static margin should be around 10%. Since our CG estimate is very crude

we assume a static margin of 15%.

Where,

From above analysis is found and hence location of wing is located.

Table 23 Location of neutral point from nose

Loaded Unloaded

4.09 2.77

Since for stability during entire mission the position of the wing body mean aerodynamic chord

should be in front of the CG hence lower value of is taken for computation. This will ensurestatic stability during the entire mission. Result of this analysis is tabulated in the table below,

Table 24 Location of leading edge from nose

Position wrt nose 2.77 2.07 1.5275


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LANDING GEAR SELECTION AND SIZING

The most commonly used arrangement for GA aircraft is the tricycle gear, with two main

wheels aft of the c.g. and an auxiliary gear forward of the c.g. With a tricycle landing gear, the c.g. is

ahead of the main wheels so that the aircraft is stable on the ground and can be landed at a fairly large

crab angle(i.e., nose is not aligned with the runway). Also, tricycle landing gear improves forwardvisibility in the ground and permits a flat cabin floor for cargo loading. Owing to these characteristics this

configuration of landing gear was selected over other discussed options.

Typically light aircraft use one wheel per strut. We have also decided to go with this option.

Wheel load geometry is shown in the figure below,

Figure 20 Wheel Load geometry

Maximum aft position of the c.g. was 3.77m, so to keep the main landing gear(MLG) behind this c.g.

location we have chosen a location 4m aft of the nose for MLG. For auxiliary landing gear a position

below the c.g. of engine was chosen which comes out to be 0.6m. This gives . Following advice fromRaymer the parameter was chosen to be 0.1 and was chosen as 0.15. The static loadson the MLG and NLG were then calculated using the following equations,

The loads are tabulated below,


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Table 25 Loads on landing gear

Load Value

Based on the above obtained data tires were selected for both MLG and NLG and the values are tabulated

below,

Table 26 Tire Data

Gear Max Loa(N) Max

Width(mm)

Max Dia

(mm)

Rolling

radius(mm)

Wheel

Diameter(mm)

Main 19,600 220 650 260 254

Nose 5,350 128 335 132 100

Recommended tire pressures for an aircraft capable of operating on tarmac runway with poor

foundation is between 345-480kPa.

STROLE DETERMINATION:

The stroke for the MLG was found from following relation,

Where,

CLOSING STATEMENTS:

The reports presents all calculations, information collected and decisions taken relating to

the design of the HR-UAS till the conceptual stage. We believe that the present design can fulfill

all the requirements that were expected of the HR-UAS. The vertical position of the C.G could

not be calculated within the given time. Hence for the same reason the landing gear could not be


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designed. The overall experience of designing this aircraft was very enriching. The design

exercise gave us a practical insight into various aspects of the aircraft designing. We would now

feel more confident while pursuing similar challenges in future.

REFERENCES:

1. Raymer, Daniel P., Aircraft Design: A Conceptual Approach, 4th Edition, AIAA, NewYork, 2006.

2. Roskam, J., Airplane Design, Roskam Aviation and Engineering Corp., Ottawa, KS,1985.

3. High Altitude Unmanned Surveillance System, Design Report, Virginia Tech4. Internet


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APPENDIX

CONSTRAINT ANALYSIS PROGRAM

tw=0.140:0.001:0.28;

landing=788*(tw./tw); %% Changed inrefined sizing

ceiling=975*(tw./tw);

loiter=2658*(tw./tw);

range=1535*(tw./tw);

pw=tw*0.454*9.81*(70/0.8)/746;

takeoff=1.5*144.3*pw*0.454*9.81/(0.3048*0.3048); %% Changed in

refined sizing

g=0.056;

%%roc1=19790*((tw-g)+sqrt((tw-g).^2-(4*0.02)/(3.14*6*0.7)));

roc2=19790*((tw-g)-sqrt((tw-g).^2-(4*0.02)/(3.14*6*0.7)));

figure

h=gcf;

plot(tw,landing,'-r','LineWidth',2)

hold on

plot(tw,ceiling,'-b')

hold on

plot(tw,loiter,'-b')

hold on

plot(tw,range,'-b')

hold on

plot(tw,takeoff,'-b')

plot(0.154,780,'*')

hold on

% plot(tw,roc1,'-r')

% hold on

plot(tw,roc2,'-r','LineWidth',2)

text(0.26,1480,'\downarrow Takeoff')

text(0.22,1595,'\downarrow Range')

text(0.23,2715,'\downarrow Loiter')text(0.23,1025,'\downarrow Ceiling')

text(0.24,850,'\downarrow Landing')

text(0.173,749.8,'\leftarrow ROC')

hold on

hr uas report

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