Aerospace Vehicle Design Design of Tactical Reconnaissance UAV (Report)

SubmittedTo:

Mr. Sher Afgan

Submitted by:

Irfan Zafar (Aero-06)

INSTITUTE OF SPACE TECHNOLOGY

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Table of Contents 1. Unmanned Aerial Vehicles (UAVs) ................................................................................................... 3

2. Request for Proposal........................................................................................................................ 7

3. Mission Profile………………………………….…………………………………………………………………………………………7

4. Reference Aircraft ............................................................................................................................ 8

5. Weight Estimation ........................................................................................................................... 9

6. Airfoil and Configuration Selection ................................................................................................ 10

7. Thrust-to-Weight Ratio and Wing Loading .................................................................................... 11

8. Second Weight Estimation ............................................................................................................. 12

9. Initial Sizing……………………………………………………………………………………………………………………………….13

10. Parachute landing System…………………………………………………………………………………………………..…….16

11. Lofting ............................................................................................................................................ 17

12. Aerodynamics ................................................................................................................................ 18

13. Structural Analysis……………………………………………………………………………………………………….……………20

14. Group Weight Estimation ............................................................... Error! Bookmark not defined.2

15. Sizing Design Freeze ........................................................................ Error! Bookmark not defined.3

16. CG variation ................................................................................................................................. 244

17. Stability and Control Analysis……………………………………………………………………………………………………25

18. Performance and Analysis ................................................................ Error! Bookmark not defined.

19. Energy Methods of Analysis .............................................................. Error! Bookmark not defined.

20. Cost Analysis…………………………………………………………………………………………………………………………..…39

21. Conclusion ...................................................................................................................................... 40

22. References ................................................................................................................................. 4041

23. Software Used………………………………………………………………………………………………………………………..…42

24. Appendix A……………………………………………………………………………………………………………………………….43

25. Appendix B………………………………………………………………………………………………………………………………..45

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1- Unmanned Aerial Vehicles (UAVs): Unmanned Aerial Vehicles (UAVs) are becoming very important, both technologically and as

business sector, with in the aerospace industry.

Unmanned Aerial Vehicles (UAVs) are making significant contributions to the war fighting

capability of operational forces. They greatly improve the timeliness of battlefield information

while reducing the risk of capture or loss of manned RECCE assets. When compared to manned

RECCE aircraft they are cost effective and versatile systems. While reconnaissance, intelligence,

surveillance, and target acquisition (RISTA) are the premier missions of UAVs, they can also provide

substantial support to intelligence preparation of the battlefield (IPB), situation development,

battle management (BM), battle damage assessment (BDA), and even rear area security (RAS) to

monitor our OPSEC posture.

HISTORY

Although much of the technology and equipment associated with the UAV are relatively

new, the concept is old. Before the US entered World War I, the US Navy (USN) developed a

seaplane that could operate without a pilot onboard. Experimentation continued on the concept

through the 1920s and 1930s. The Navy also developed and used small plywood UAV in the Pacific

in World War II to attack heavily defended targets.

The Army Air Corps experienced heavy losses of aircraft and trained aircrews in World War

II and thus the Aphrodite Project was conceived. Aphrodite used old B-17 aircraft loaded with

explosives, flown to altitude by a pilot, who then bailed out. A second B-17 assumed radio control of

the unmanned aircraft and directed it to crash into a target. After World War II, drone B-1 7's were

used in atom bomb tests in the South Pacific.

At Fort Huachuca in the late 1950s the Army placed cameras on target drones and

developed an operational UAV reconnaissance system. Several years later they replaced the camera

with a television system.

By 1964, an Air Force drone reconnaissance program, known as Buffalo Hunter, was under

full development. A C-1 30D aircraft could carry up to four drones under its wings, flying out of

Vietnam they would launch them like missiles on a preprogrammed flight over enemy held

territory. From the mid-1960s until the end of the Vietnam War, more than 3,000 missions were

flown over North Vietnam and China.

The Navy also used UAVs during the Vietnam War. One program, called DASH, was a remote

helicopter carrying a television camera and two 250-pound torpedoes was used to detect and

destroy North Vietnam supply barges in Mekong Delta waterways. Although this program enjoyed

several successful missions, the helicopter flight gyroscopes were not up to standard and the

program was discontinued.

By the mid-1980s, a joint project, the Pioneer UAV system, came into being. The Pioneer

was used in Operations Desert Storm and provided outstanding intelligence and fire support

information to the commander.

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Other countries such as Israel, the former Soviet Union, and several European countries

have also integrated UAVs into military operations with a high degree of success.

During Operation Iraqi Freedom, the need for timely and actionable intelligence underlined

the increased need for UAVs. The Chicago Tribune reported in a November 23, 2004 story, that

Army Brigades operating in Iraq had the number of UAVs assigned to them increase from two to

three. Similarly, each brigade was to be assigned four teams of 22 soldiers to operate the 450

unmanned aerial vehicles reported to be in Iraq.

UAV Classes

UAVs are divided by class category. During operations where more than one system is available,

UAV systems can be task organized and class categories selected to achieve the required flexibility

and capability. Listing of UAV class categories as recognized by the Department of Defense (DOD)

UAV Master Plan are given below. The categories are by range or flight hours, or both.

UAV-Close Range (UAV-CR): Operational range will be approximately 50 kilometers

UAV-Short Range (UAV-SR): Flight duration of 8 to 10 hours designed to penetrate into

enemy airspace out to a range of 200 kilometers with data link

UAV-Endurance (UAV-E): Minimum of 24 hour's coverage and be capable of performing

multiple missions simultaneously

While there are differences in range and capabilities, all of these categories of UAVs are

considered to be members of the family of UAVs. The family of UAVs concept is based upon

commonality and interpretability. All ground receivers are capable of receiving the video of any

other UAV within range, regardless of the class category.

Unmanned Aerial Vehicles (UAVs) are remotely piloted or self-piloted aircraft that can carry

cameras, sensors, communications equipment or other payloads. They have been used in a

reconnaissance and intelligence-gathering role since the 1950s, and more challenging roles are

envisioned, including combat missions. By the early 1990s UAVs were sought to satisfy surveillance

requirements in Close Range, Short Range or Endurance categories. Close Range was defined to be

within 50 kilometers, Short Range was defined as within 200 kilometers and Endurance as anything

beyond. By the late 1990s, the Close and Short Range categories were combined, and a separate

Shipboard category emerged. The current classes of these vehicles are the Tactical UAV and the

Endurance category.

Pioneer: Procured beginning in 1985 as an interim UAV capability to provide imagery intelligence

for tactical commanders on land and see at ranges out to 185 kilometers. No longer in the Army

inventory (returned to the US Navy in 1995).

Tactical UAV: Designed to support tactical commanders with near-real-time imagery intelligence

at ranges up to 200 kilometers. Outrider Advanced Concept Technology Demonstration (ACTD)

program terminated. Material solution for TUAV requirements is being pursued through a

completive acquisition process with goal of contract award in DEC 99.

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Joint Tactical UAV (Hunter): Developed to provide ground and maritime forces with near-real-

time imagery intelligence at ranges up to 200 kilometers; extensible to 300+ kilometers by using

another Hunter UAV as an airborne relay. Training base located at Fort Huachuca, with additional

baseline at Fort Polk to support JRTC rotations. Operational assets based at Fort Hood (currently

supporting the KFOR in Kosovo).

Medium Altitude Endurance UAV (Predator): Advanced Concept Technology Demonstration

now transitioned to Low-Rate Initial Production (LRIP). Provides imagery intelligence to satisfy

Joint Task Force and Theater Commanders at ranges out to 500 nautical miles. No longer in the

Army inventory (transferred to the US Air Force in 1996).

High Altitude Endurance UAV (Global Hawk): Intended for missions requiring long-range

deployment and wide-area surveillance (EO/IR and SAR) or long sensor dwell over the target area.

Directly deployable from CONUS to the theater of operations. Advanced Concept Technology

Demonstration (ACTD) managed by the US Air Force.

Tactical Control Station (TCS): The Tactical Control Station is the software and communications

links required to control the TUAV, MAE-UAV, and other future tactical UAV's. It also provides

connectivity to other C4I systems.

Micro Unmanned Aerial Vehicles (MAV): DARPA program to explore the military relevance of

Micro Air Vehicles for future military operations, and to develop and demonstrate flight enabling

technologies for very small aircraft (less than 15cm/6in. in any dimension).

The Air Forces are interested in aggressively pursuing emerging technologies to develop,

field, and operate UAV solutions for applicable military roles across the spectrum of warfare to

meet validated needs within specific mission areas based on cost, capability, reliability and

suitability. Consistent with these objectives, the air vehicles directorate is interested in proposals

for technology development which demonstrate revolutionary potential to cut cost, weight or

increase performance. Candidate technologies include: Automated Flight Control Systems,

Autonomous Aerial Refueling, Flexible Structures, Low-cost Composite Technology, Multifunctional

Integrated Structures and Subsystems, Active Flow Control as well as Preliminary Design Tools for

Advanced Technologies

Uninhabited Air Vehicles have many commercial applications. Mail delivery, automated

pipeline monitoring, drug enforcement and border control are but a few examples of the potential

commercial uses. In addition, the revolutionary technologies of interest to the Air Vehicles

Directorate also have dual use potential. Automated flight control has application to military and

commercial UAVs in operations in mixed manned and unmanned air space and to general aviation

in terms of reducing pilot work. Autonomous refueling can reduce control-related accidents as well

as greatly enhance UAV endurance and range. Flexible and adaptive structures technology can

extend the range and reduce maintenance costs for military and commercial high speed transports.

Also, increased application of microprocessor technology in smart commodity products will also

benefit from this technology by allowing embedding of electronic chips, wiring, power and other

devices. Low Cost Composite Technologies and multifunctional integrated Structures and

Subsystems have a wide variety of applications in commercial aviation and transportation vehicles

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where cost and weight are extremely critical for commercial viability. Active flow control

technology can reduce drag and wheel well noise on military and commercial transports.

Preliminary design tools can increase the effectiveness of all these technologies as these

technologies come together in military and commercial products.

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2- Request Proposal: Aim is to design an autopilot with radio control backup with a primary mission to fly over

predefined targets for reconnaissance purpose remaining undetected by radars of enemy and land

back. The tactical reconnaissance UAV is usually designed not to loiter over the target area, and

real-time intelligence is less essential.

Requirements:

1. Should have a range of 1200nm (2222 km). 2. Specified mission relevant payload is 520 lb. 3. Capability of LRV (Launch Recovery Vehicle) launch. 4. Specified cruise conditions of Mach No. 0.8 at an altitude of 20,000ft. 5. Should be capable of achieving an altitude of 43,000ft. 6. Shouldn’t necessarily possess capabilities of loitering and instantaneous turn.

3- Mission Profile:

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4- Reference Aircraft: Different UAVs were compared by rating them with respect to their similarities with the specified requirement of the UAV being designed.

M324, an Egyptian UAV, was selected as a reference aircraft. Specification for the reference aircraft are:

Length 20.17 ft Wingspan 10.95 ft Height 2.83 ft Max Gross Weight 2,384 lb Speed (Max) Mach 0.85 Max Altitude 43,000 ft Cruise Range 1,380 Nautical Miles Engine TCAE 373-8 Turbojet Engine

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5- First Weight Estimation: Selection of Mission segment weight fractions is a step in design as suggested by Raymer.

These calculations are based upon fixed values of weight fractions for some mission segments and Range and Endurance concerns during Cruise and Loiter conditions, respectively. Certain Trade-off studies were also done which include Range, Payload and Composite Materials Trade. The mission requirements for which the weight estimation was done are given as under:

Wo SFC Range L/D Payload V W1/Wo W2/W1 W3/W2 W3/Wo Wf/Wo We/Wo Wo

Climb Cruise Descent

3000 0.00025 7310526 12.1 520 830 0.985 0.83362 0.9917 0.8143 0.19684 0.48263 1622.3

1622.3 0.00025 7310526 12.1 520 830 0.985 0.83362 0.9917 0.8143 0.19684 0.50991 1773.3

1773.3 0.00025 7310526 12.1 520 830 0.985 0.83362 0.9917 0.8143 0.19684 0.50673 1754.2

1754.2 0.00025 7310526 12.1 520 830 0.985 0.83362 0.9917 0.8143 0.19684 0.50713 1756.6

1756.6 0.00025 7310526 12.1 520 830 0.985 0.83362 0.9917 0.8143 0.19684 0.50708 1756.3

(

) (

)

( ) ( )

Wo = 1756 lb

Trade-off Studies:

Varying the range and payload values, trade off was studied between the varying the values and the respected calculated weight.

1600

1650

1700

1750

1800

1850

1900

1950

2000

800 1000 1200 1400 1600

Range variation

1000

1500

2000

2500

3000

200 400 600 800

Payload Variation

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6- Airfoil and Configuration Selection:

Airfoil Selection: Plane being designed has very simple mission profile and does not require too many

maneuvers. So the first preference was to reduce drag for cruise portion of flight as most of the flight duration involves cruise. Wind tunnel and flight tests of 6-series airfoil showed extensive laminar boundary layers also at comparatively larger values of Reynolds number as compared to other airfoils. 6-series airfoil also gives extremely low drag coefficient near the design lift coefficient.

From the historical trend curve given in the Raymer in fig. 4.14: for design Mach number 0.8

we have thickness ratio (t/c) as:

t/c = 15 %

So from NACA sixth series, for the obtained thickness ratio and from the conventional data

of the other planes, airfoil selected is NACA 66-415. Shape of the selected airfoil is shown below:

Airfoil Cl-α , Cd-α, Cl/Cd-α and Cm-α curves are given in the Appendix A. For horizontal and

vertical tail, similar airfoil was used.

Wing Configuration: A mid-wing configuration was selected to minimize drag. Straight wings were selected to

maximize lift at the operational flight envelope, which is low subsonic. Tapering was selected in

order to minimize drag due to lift and make the lift distribution closer to elliptical.

After studying the effects of effect of wing planform and setting drag reduction during low

subsonic cruise as the first priority, a straight, tapered and high aspect ratio configuration for the

wing was selected. Hoener tips were selected to minimize the downwash effect.

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Tail Configuration: For parachute landing mechanism, it is mandatory to select a suitable place for parachute storage so that after being ejected, parachute does not damage aircraft surfaces around its ejection point. For this reason and some study for different types of tails used, suitable tail configuration selected was twin tail configuration as shown below. Parachute will be stored between the two vertical tails.

7- Thrust-to-Weight Ratio and Wing Loading:

Thrust to weight ratio:

For cruise,

= 7.9672, from the equation given in Raymer:

= 0.126

For take-off, from the parameters and equation given in Raymer:

Wing loading:

For landing, Vstall = 90 knots = 151.9 ft/s, CLMAX = 1.12, ρ = 0.0023769 slugs/ft3 at sea level. Using

( ) )

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For cruise conditions; CD0 = 0.015, Vmax = 830 ft/s, ρ = 0.0012673 slugs/ft3 at 20,000 ft:

For ceiling; ρ = 0.0005087 slugs/ft3 at 20,000 ft, Vmax = 774.36 ft/s:

8- Second Weight Estimation:

Second weight estimation was done using more accurate procedure given in the book. Some

of the data used for the weight estimation is listed below in the table:

Wo SFC Range L/D Payload V We/Wo W1/Wo W2/W1 W3/W2 W3/Wo Wf/Wo Wo

3000 0.000269 7310526 12.3 520 830 0.544 0.9805 0.8248 0.99 0.8006 0.21134 2125

2125 0.000269 7310526 12.3 520 830 0.4887 0.9805 0.8248 0.99 0.8006 0.21134 1733.8

1733.8 0.000269 7310526 12.3 520 830 0.4792 0.9805 0.8248 0.99 0.8006 0.21134 1680.6

1680.6 0.000269 7310526 12.3 520 830 0.4787 0.9805 0.8248 0.99 0.8006 0.21134 1677.5

1677.5 0.000269 7310526 12.3 520 830 0.4786 0.9805 0.8248 0.99 0.8006 0.21134 1677.3

Wo = 1677.3 lb

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9- Initial Sizing: Fuselage Length:

From table 6.3 in Raymer, calculated fuselage length is:

Fuselage Length = 17.35 ft

Wing Sizing: Based on the equations provided in the Raymer’s book, the following wing planform was

designed.

Wing Geometry

Tip to chord ratio (t/c) 15% Ref. area (Sref) 59.1 ft2 Aspect ratio (A) 3.6 Wingspan (b) 14.59 ft Leading edge sweep(ΛLE ) 30: Quarter chord sweep (Λ c/4 ) 27: Taper ratio (λ) 0.608 Root chord (Croot) 4.2 ft Tip chord (Ctip) 2.55 ft

Three views for the wing planform were plotted on AutoCAD and are shown:

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Tail Sizing:

From the procedure and historical trend values provided in the Raymer, calculated parameters for tail are:

Horizontal Tail

Cht 0.40

b 5.27 ft

S 7.73 ft2

Vertical Tail

Cvt 0.07

b 6.24 ft

S 10.82 ft2

Control Surfaces’ Sizing:

Final sizing of these surfaces is based upon dynamic analysis of control

effectiveness, including structural bending and control systems effects. For initial design,

historical guidelines are used.

Ailerons:

Ailerons typically extend from about 50% to 90% of the span and about 15% to 25% of the

wing chord. For this specified aircraft, ailerons are placed 0.7295 ft from wing tip and 3.6475 ft

from wing root. For 70% span of wing:

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Elevators:

Elevators typically extend to the tip of the tail or to about 90% of the tail span and about 25% to

50% of the tail chord. For 90 % span

Rudders:

Rudders typically extend to the tip of the tail or to about 90% of the tail span and about 25% to

50% of the tail chord.

Propulsion system:

Developing altogether a new engine costs several billion dollars, so most aircraft development projects do not rate the development of development of new engine and rely on selecting the best of the existing engines. So was in our case. From the reference plane, jet engine being assumed for the design plane is General Electric (USA)-J85-17.

Some performance improvements due to the use of newer technologies will be made. Scaled engine parameters come out to be:

L = Lactual(SF )0.4 = 2.15 ft

D = Dactual(SF )0.5 = .8305 ft W= Wactual(SF )1.1 = 113.04 lb

After applying 10% improvements due to modern technologies, following two parameters changes:

SFC = 0.873 lb/h = 0.000243 1/s W = 101.736 lb

Inlet is placed over the fuselage for the comfort of LRV launch mechanism and also due the

reason that we may not be catering higher angles of attack that may cause the flow to separate from the rare part of fuselage. Pitot inlet is found to be suitable as for subsonic speed, pitot type will have virtually 100% pressure recovery VS about 90% for well designed NACA inlet. Capture Area Calculation:

From literature given in the book, mass flow can be estimated as:

Mass Flow = 17.93 lb/s

For M = 0.8, from fig. 10.16:

Acapture = 0.44825 ft2

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Fuel systems

The fuel weight to be carried is 362 lb. They type of fuel to be used is Aviation Gasoline. The volume

required for its storage is calculated as follows:

Fuel systems

Fuel density (lb/gal) at MIL spec

6

Fuel volume (gal) 60.34

Fuel volume (cubic ft) (7.5 gal = 1 cubic ft)

8.05

Fuel will be stored in integral tanks in the fuselage.

10- Parachute Landing System: In UAV, a parachute landing system is being used for causing the aircraft to land while being

suspended by a parachute. Parachute is stored in the fuselage compartment that is open able and pivot able in outward direction. A particular parachute will be arranged in the fuselage of the aircraft that will be expelled in the beginning of the landing phase so that the craft can drop to ground in a horizontal position and at a speed that permits safe and soft landing.

This type of landing of a craft will be, however, rather difficult to realize, particularly if the craft has to land in a precisely defined area. In other words, the localization of the landing process will be difficult to achieve.

Flaps are arranged such that being pivotal in outward direction about an axis about which flap can be locked into the pivoted position to provide a protection and shielding wall between the tail section of the aircraft and the deploying parachute. All these features are deemed to render the system economical but also reliable under consideration of all the various requirements delineated above.

11- Payload:

Raytheon Space & Airborne Systems supplies integrated sensor suite (ISS) which includes the

synthetic aperture radar and the electro-optical and third-generation infrared sensor system. This system

will be used in the UAV.

A 10in reflecting telescope provides common optics for infrared and electro-optical sensors. The

electro-optical / infrared sensor operates in the 0.4 to 0.8 micron visible waveband and the 3.6 to 5-micron

infrared band. In spot collection mode the coverage is 1,900 spots a day with spot size 2km² to a

geological accuracy of 20m circular error of probability. In wide area search mode, the swath is 10km

wide and the coverage is 40,000nm² a day.

The synthetic aperture radar and ground moving target indicator (GMTI) operates at X-band with

a 600MHz bandwidth, and 3.5kW peak power. The system can obtain images with 3ft resolution in its

wide area search mode and 1ft resolution in its spot mode.

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Raytheon is contracted to supply one enhanced integrated sensor suite (EISS) which is said to

improve the range of both SAR and infrared system by 50%.

The Raytheon ground station receives the high-quality imagery obtained by the air vehicle sensor

suite. The ground system forwards the imagery to military commanders and users in the field.

Other payloads include hyper spectral sensors for chemical and biological agent detection.

12- Lofting:

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13- AERODYNAMICS:

Lift Curve Slope:

Lift curve slope is needed during conceptual design for various reasons. First it is usually used to

set the wing incidence angle. Also the wing incidence angle influences the the required fuselage

angle of attack during take-off and landing. Also the calculations of drag due to lift uses the slope of

the lift curve. Also we have to determine lift slope curve for longitudinal stability analysis. Lift curve

slope for subsonic regime is given as;

Where ‘F’ is the fuselage lift factor given as,

( ⁄ ) And

, So,

( )

So, =0.6

( )

Maximum Lift Coefficient:

Estimation of maximum lift is probably the least reliable of all calculations used in the

aircraft conceptual design. Even refined wind tunnel tests cannot predict maximum lift with great

accuracy. Maximum lift coefficient of the clean wing will usually be about 90% of the airfoil’s

maximum lift as determined from the 2-D airfoil data at a similar Reynolds number. So the

expression is given as:

cos

=27O

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Subsequently, the angle of attack for maximum lift, as expressed in the equations and figures given

in literature, is estimated as:

αCLmax = 13.5 deg

Drag polar:

CD = CD,e + CD,i

From the procedure and formulae given in Raymer:

CD = CD0 + KCL2

CD0 = 0.017

K = 0.0957

Calculation of the above obtained values is attached. So drag polar equation comes out to be:

CD = 0.017 + 0.0957 CL2

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14- Structural Loads:

V-n Diagram:

V-n diagram was plotted in AAA software.

Material Selection: A number of properties are important to the selection of materials for aircrafts. Selection of

best material depends upon the application. Factors to be considered include yield and ultimate strength, stiffness, density, fracture, toughness, fatigue and crack resistance, creep, corrosion resistance, temperature limits, producibility, repair ability, cost and availability. Composite materials have revolutionized aircraft materials industry and have helped by reducing the structure weights up to 25%.

The choice of composite materials in aircraft structures was dictated by a need to reduce weight and to improve strength, reliability and maintainability. As shown in the figure, the composite fraction of the structural weight for fighter and attack aircraft seems to be leveling off at 30 percent. Similarly relative ratio of the composite materials used for different category of airplanes is shown below in the graph. Based on the prevalent applications seen by composites in different aircrafts, this limit is an indicator of lack of confidence in composite applications in highly 3-D loaded fuselage and wing substructures such as the main spars, and bulkheads.

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From the plane being designed, design requirements give clue of fall of the UAV in both transport and military purposes. UAV should have optimum reconnaissance capabilities as well as able to carry specified payload. So both categories were studied to know about the materials being used in their structures.

The Euro fighter Typhoon uses composite materials in the wing skins, forward fuselage, flaperons and the rudder. Euro fighter’s exterior skins are made of Hexcel’s 8552 toughened epoxy and constitute 70% of the wetted area. Overall, 40% of the Euro fighter’s structural weight consists of carbon-fiber composite materials.

For the B-2, stealth or minimizing the radar cross-section was the primary driver and as

such carbon fiber composites were extensively used in the primary structure to offset the weight penalty from radar absorbing materials applied to the exterior.

Significant use of composites in commercial transports has been on the Boeing 777.

Composite structures make up 10 percent of the structural weight of the B-777. Figure shows the various composite structural elements used in the B-777. The composite empennage alone saves approximately 1500 lb over similar aluminum structure. In the case of transport aircraft where cost and reliability are the predominant factors, composite applications seem to be leveling off at 20 percent of the structural weight a ceiling lower than for combat aircraft. The barrier in this case is set by the affordability of the airframe since initial acquisition cost plays a major role in airlines’ selection of a particular model.

Based on the discussion above and also from the helping material given in the Raymer’s

book, below given is rough idea of the different material being used in different parts of UAV:

Aircraft Section Material used

Basic Structure that includes fuselage, wings and tails substructures

Steel and its alloys

Wing and Tail skins Carbon fibers, such as Hexcel’s IM7, with improved strength and stiffness properties

Centre and aft Fuselage skin all-carbon/toughened- epoxy

Frontal Fuselage skin Aluminum

Firewalls and Engine mounts Steel and its alloys

Joints and fittings between structural components

Titanium

Parachute storage doors

Graphite-aramid-epoxy hybrid composite

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15- Group Weight Estimation:

Third weight estimation was carried out using the equations given in Raymer literature.

Following values have been estimated and fudge factors were applied for the weight compensation

due to the use of composite materials and modern material technologies.

Section Weight Estimated

Weight (lb)

Fudge Factor

Adjusted Weight

(lb)

W wing 92.28 0.85 78.4 W Horizontal Tail 13.6 0.83 11.3

W Vertical Tail 12.71 0.83 10.6 W Fuselage 117.95 0.90 106.2

W Installed Engine 182.67 182.7 W Fuel System 35.18 35.2

W Flight Controls 12.41 12.4

W Hydraulics 0.202 0.20 W Electrical 165.28 165.3

W Avionics 121.1 121.1 W Furnishing 32.20 32.2

Total Empty Weight 755.6 Fuel Weight 437.6

Mission Relevant Payload Weight 520

Total Weight (lb) 1713.2

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16- Sizing Design Freeze The sizing obtained in the previous section was the result of several iterations, going back and forth

between different sections. The most significant change was the revision of the TOGW from 1677 lb

to 1756 lb. No more sizing will be changed from this stage onwards.

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17- CG Variation: During flight, only fuel is being. Fuel tank is placed in the fuselage at the centre. As the fuel is

consumed, c.g. will vary only in z-direction. C.g. will not vary in x-direction. In steady level condition, c.g.

is at

X = 7.8 from nose

18- Stability and Control Analysis:

Stability derivatives were calculated using DATCOM. Input and output files are shown in Appendix B.

Longitudnal Stability Derivatives:

Using DATCOM and Blakelock, the stability derivatives are computed as follows:

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Solutions of Longitudinal EOM:

0)1477.0872.0)(0114.00546.0( 22 SSSS

The roots are

iS 1034.00273.0 and 2299.0S ; 6422.0S

Which give

Phugoid Mode Short Mode

t1/2=25.26 seconds and t1/2=1.58 seconds

Lateral Stability Derivatives

Similarly for Lateral stability derivatives:

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Solution of Lateral Equations 2( 0.5415 0.3483)( 0.02334)( 0.6121) 0S S S S S

This gives the following solution

Roll subsidence

TR=1/0.6121=1.633 seconds

Spiral Divergence

TS=1/0.02334=42.8449 seconds

Dutch Roll

Frequency=0.5902 rad/s

Damping Ratio=0.4618

T0.5=2.54 seconds

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19- Performance Analysis:

Aerodynamic Analysis:

For Mach no. = 0.8; Drag polar and lift to drag ratio curve is given as:

Also the tangent line to the drag polar from the origin locates the point of maximum lift-to-drag ratio for the airplane. This point is shown above.

From graphs, the point comes out to for:

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CL = 0.4

(L/D) max = 12.3

This point is sometimes known as the design point for the airplane and the corresponding lift coefficient is known as design lift coefficient.

Propulsion Analysis:

Variation with Altitude and Velocity:

Generally in the analysis of airplane performance in the cruise range, it appears reasonable to assume thrust constant.

Also for turbojet,

(

)

Plotting for the above relation, we get:

Thrust Specific Fuel Consumption is also assumed to be constant with varying altitude and mach numbers due to negligible variation.

Power Available versus Power required:

Turbojets are rated in terms of thrust. So for power available:

PA = TAV∞

0

100

200

300

400

500

600

700

800

900

0 10000 20000 30000 40000 50000

Thru

st A

vaila

ble

Altitude

Thrust variation with Altitude

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And for power required:

PR = TRV∞

Relation between both and their variation with velocity is shown in the graph as:

Note that the two curves meet again to give a point representing maximum velocity for the airplane. As specified, maximum velocity comes out to be 830 ft/s.

Steady-Level Flight Performance analysis:

For steady and level flight, equations of motion are:

Lift is equal to weight and trust required for non-accelerated level flight is equal to drag. For thrust required (drag), we have:

(

) ( ) (

) ( )

In the above expression, first term on the right side represents zero lift drag while second term represents drag due to lift. Plotting three curves for thrust required, zero lift drag and drag due to lift respectively for variation with velocity, we obtain the following graph:

0

100000

200000

300000

400000

500000

600000

700000

0 200 400 600 800 1000 1200

Po

we

r

Velocity (ft/s)

Maximum Velocity

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Straight black line shown in the above figure differentiates the regions of velocity stability and instability. On the right side of the line, region is of velocity stability and vice versa. Also the upper curve represents the total drag curve. Corresponding velocity gives the velocity for the minimum thrust required and maximum lift to drag ratio.

Climb Performance:

Rate of climb for any airplane is given analytically as:

V∞Sinθ = R/C = ( ) ( )

As we have plotted the power available and power required curve already as:

0

100000

200000

300000

400000

500000

600000

700000

0 200 400 600 800 1000 1200

Po

we

r

Velocity (ft/s)

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The difference between the two curves gives the value of excess power (TV∞-DV∞) at the corresponding point, we can obtain maximum rate of climb where this difference is maximum. Analytically, we obtain from the expression:

(R/C) max =215.24 ft/s

Also we find velocity corresponding to the (R/C)max :

V(R/C)max = 674.82ft/s

Graph elaborates the analytical calculated values as the maximum excess power occurs at the velocity of the maximum rate of climb.

Also we find θmax and Vθmax analytically as:

Θmax = 24.6⁰

Vθmax = 132.4ft/s

We also draw Hodograph diagram for climb performance at sea level:

From graph, it is obvious that the analytical quite match with the graphical values.

Gliding Performance:

For analytical values:

Vv =

√

(

(

)

)

(

) = 28.5 ft/s

0

20

40

60

80

100

0 200 400 600 800 1000

Rat

e o

f C

limb

(ft

.lb

/s)

V horizontal

Hodograph

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Θmin =6.5⁰

Range:

Analytical value is obtained to be:

R = (

) (√

)(

)(( (

)) (

)) =

Maximum range is obtained as:

Rmax = 1120 nm

Endurance:

Analytical value for endurance can be found as:

E = (

) (

) (

)

Maximum endurance is obtained as:

Emax = 3 hrs 20 mins.

-250

-200

-150

-100

-50

00 500 1000

Rat

e o

f d

esc

en

t

V horizontal

Hodograph for unpowered flight

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20- Energy Method of Analysis:

Specific Power vs. Velocity graph:

0

20

40

60

80

100

120

140

160

180

0 0.2 0.4 0.6 0.8 1

Ps

Mach No.

h=0

h=5000

h=10000

h=15000

h=20000

h=30000

h=35000

h=40000

h=25000

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Altitude-Mach No. map:

0

10000

20000

30000

40000

50000

60000

70000

0 0.2 0.4 0.6 0.8 1 1.2

Alt

itu

de

Mach No.

Altitude - Mach No. map

Ps = 60Ps = 80Ps=100Ps=120Ps = 140Ps=40Ps=20

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Maximum Rate of Climb Schedule:

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Optimum Energy Climb Schedule:

Minimum Time to Climb:

Minimum Time to Climb

he dhe/dt 1/(dhe/dt)

24000 120 0.00833

30000 100 0.010

36000 80 0.0125

43000 60 0.0167

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By calculating the area under the curve, we can find the minimum time to climb for specific altitude. In this case for climbing from 24,000 ft to 43,000, aircraft will take 238 seconds or 4 min approximately.

Minimum fuel to climb:

Minimum Fuel to Climb

he (dWf/dt) dhe/dt (dhe/dWf) 1/(dhe/dWf)

24000 0.096 120 1250 0.0008

30000 0.0775 100 1290.3 0.000775

36000 0.0537 80 1490 0.000671

43000 0.0443 60 1354.4 0.000738

0

0.002

0.004

0.006

0.008

0.01

0.012

0.014

0.016

0.018

20000 25000 30000 35000 40000 45000

1/(

dh

e/d

t)

he (ft)

Minimum time to climb

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By calculating the area under the curve, we can find the minimum fuel to climb for specific altitude. In this case for climbing from 24,000 ft to 43,000, aircraft will consume 13.76 lb approximately.

0.00066

0.00068

0.0007

0.00072

0.00074

0.00076

0.00078

0.0008

0.00082

20000 25000 30000 35000 40000 45000

1/(

dh

e/d

Wf)

he

Minimum fuel to climb

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21- Cost Analysis:

Cost analysis was done for 10 UAV productions in 5 years from the formulae given in Raymer. Values come out to be:

Engineering Cost 42,127,530 $ Tooling Cost 19,327,616 $ Manufacturing Cost 16,940,453 $ Devel Support Cost 20,433,218 $ Flight Test Cost 11,995,674 $ Manufacturing Cost 3,516,130 $ Engg. Production Cost 39,657,166 $ Total Cost 154 million dollars

So,

Estimated Cost for one UAV = 15.4 million dollars

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22- Conclusion:

The 1200 nm range was almost achieved, lesser than reference aircraft. This can be attributed to the decrease in take-off gross weight from 2300lb lb to 1700 lb, without any increase in the fuel ratio to be carried. Note that the 2300 lb M324 carries fuel which 27.2% of its weight. The designed aircraft carries only 21.3 % fuel by weight

The objective of having an operational altitude of 20,000 ft was achieved successfully. The absolute ceiling of the aircraft (43,000 ft) is same as that of the reference aircraft M324. The objective of carrying 520 lb internal payload was also achieved. Had there been same payload of 320 lb, weight would have been 1502. A payload weight increase of 63.3% increased the take-off gross weight by 14.05%.

The maximum range of the aircraft came out to be approximately 1200 nautical miles (2400 km) giving it an operational radius of 600 nautical miles (1200 km). This means that in theory, this aircraft, taking off from Islamabad could perform missions directed at targets in Peshawar (141 km), Karachi (1142 km), Srinagar (159 km), New Delhi (682 km), Fatehpur (1132 km), Kabul (378 km), and Kandahar (732 km).

The stability analysis was carried out at the gear-up configuration immediately after take-off. The c.g. plot shows a rearward shift of cg which needs to be compensated for through further analyses. For now, the longitudinal stability is satisfactory at best. The lateral directional stability is also marginally satisfactory.

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References

Raymer, Daniel P., Aircraft Design: A Conceptual Approach, Fourth Edition

Lang, James D., Aircraft Performance, Stability and Control, Vol I

Blakelock, John H., Automatic Control of Aircraft and Missiles

Cook, M.V., Flight Dynamics and Principles

Abbott, Theory of Wing Sections

Users Manual, THE USAF STABILITY AND CONTROL DATCOM

General Atomics Aeronautical Systems Inc. http://www.ga-asi.com/

Federal Aviation Regulations: FARs Part 23, Flightsim Aviation Zone

http://www.flightsimaviation.com/data/FARS/part_23-appA.html

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Software used

MS Excel

MS Paint

AutoCAD

USAF Digital Datcom

USAFA AeroDYNAMIC

DARCorporation Advanced Aircraft Analysis

MATLAB

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Appendix A:

Airfoil Curves:

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Appendix B:

DATCOM Input File:

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DATCOM output file:

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