Design Proposal Presentation Final Slides

8/13/2019 Design Proposal Presentation Final Slides

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R YA N S T E V E N S , J A M E S M I L L A N E

PAT R I C K N O R M A N , A N D R E W C U L L

E K I N O R E R , B R I A N O N E I L L

N O L A N L A H R

AAE 251 Reconnaissance UAV

Design Proposal


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Reason for Need

Increasing amount of space debris causes a need for non-orbital

surveillance

DARPA has requested a design for a surveillance UAV in

conjunction with a rocket based launch system

12


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Rocket Requirements

The rocket-based launch system priority is to reach its destination as

quickly as possible, while being cost efficient.

Parachutes will be used to decelerate payload to desired speed

Launch system propels payload into elliptical orbit with a periapsis

located beneath Earths surface. No substantial land mass may reside within 30 degrees in either

direction of the launch inclination for 500 km.

Assume the cost of inert mass to be $500/kg for solid propellant-based

stages and $1000/kg for liquid propellant-based stages. Assume the cost of the propellants to be $20/kg for both solid and

liquid propellants.


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UAV Requirements

The UAV must be: deployed during the descent of the orbit at an altitude of 40,000 feet.

deployed at a flight velocity of 350 mph.

capable of loitering over the desired region for a minimum of 24 hours.

capable of loitering at an altitude of 50,000 feet.

capable of traveling 3,000 nautical miles to land after completing theperiod of surveillance.

capable of carrying a standard payload of surveillance equipment

capable of landing on runway or aircraft carrier

comparable to fuel mass fraction of existing UAVs

able to fit inside the rocket fairing

Assume jet fuel to be $4/kg

Assume mass to be $100/kg


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UAV Performance Parameters

Airfoil: NACA 2415 at 2: 0.45

at 2: 0.00951

Wingspan: 16

Wing Reference Area: 19.4

Aspect Ratio: 13.2

Max

(airfoil only): 61.53

Max

(fuselage included): 47.33

Endurance: 64.82

Weight (unfueled): 1700 Weight (fueled): 2900

Landing Distance: 3460

Landing Velocity: 124

Powerplant: Rolls-Royce F137-AD-100

Thrust: 8290

Cruise Speed: 372

Service Ceiling: 83000


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Launch Vehicle Combined Parameters

Total Rocket Mass: 165,040 kg

Payload Mass: 3,500 kg(includes plane and fairing)

Mission delta-V: 7,900 m/s(first stage provides 57%)


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Launch Vehicle Parameters

First Stage Total Mass: 148,813 kg Engine: RD-191 Fuel: LOX/Kerosene ISP (Sea level): 311 s ISP(vacuum): 337 s Engine Mass: 3,230 kg Engine Thrust: 2,079 kN Inert Mass: 7,813 kg Structure Mass: 4,583 kg Propellant Mass: 141,000 kg Inert Mass Fraction: 0.0525

Second Stage Total Mass: 12,724 kg Engine: RD-58M x2 Fuel: LOX/Kerosene ISP (vacuum): 353 s

Engine Mass: 230 kg x2 Engine Thrust: 83.40 kN x2 Inert Mass: 1,141 kg Structure Mass: 681 kg Propellant Mass: 11,583 kg Inert Mass Fraction: 0.0879


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Max weight of 5000 kg

Max length of 10 meters for

UAV

Max radius of 2 meters forfairing

Designed after Titan II and

Atlas G

Designed to be a cross

between satellite fairing andre-entry vehicle

Rocket Fairing

Atlas G

YRQ-0X Launch Vehicle

Fairing Titan II


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Wings

NAME CL CD CL/CD STALL

NACA 2415 0.41 0.006663 61.53 8

NACA 4415 1.643 0.029657 55.4 14

NACA 1412 1.098 0.023512 46.7 7

NACA 23012 1.095 0.025644 42.7 8.5

WORTMANN FX-72MS-150B 2.116 0.054341 38.939 11

NACA 1408 0.852 0.023342 36.5 3.5HQ 0/7 0.475 0.016102 29.5 3

To maximize endurance, maximize Cl/Cd To maximize range, minimize Cd

*Note: Adjusted Cl/Cd

by dividing by 4/3.

The final lift to drag is

47.33.

=1

ln

= 2

2

1

( )


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Fuselage

The main concerns for the design of the fuselage were

size and weight.

The total length of the aircraft could be no longer than 10

meters We decided to go with a length of 9 meters to leave room for

parachutes

The fuselage could be no wider than 3.8 meters to fit

inside rocket fairing. We ended up setting the maximum width of the fuselage to 1.8

meters to give ample room for the folding wings.


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Fuselage(cont.)

We decide to make the fuselage out of lightweight

composites to minimize weight.

Because the engine greatly affects to the center of gravity,

the sensor package was placed at the front of the plane tomove the CG as far forward as possible

The engine was placed inside of the plane to reduce drag.


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Horizontal Stabilizers

Surface area of the Horizontal Stabilizer isapproximated by:

= .15 =2.91

AR= 4.5 Span:2.8m.


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Vertical Stabilizer

Surface area of the Vertical Stabilizer isapproximated by:

= 0.09 = 1.746 = 0.873 (each)

AR = 0.9 Span = 0.886


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Powerplant

Low TSFC for optimalendurance

Allison F137-AD-100

8290 lbf .39 TSFC

=

/


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Range and Endurance

We needed 40 hours of endurance to ensure abilityto meet 24 hour requirement and flight time to base. Weight of the:

Structure:800 kg

Equipment: 200 kg Turbofan: 700 kg

Fuel: 1200kg

Lift to Drag Ratio 47.33

TSFC 0.39

Empty Weight 3746.8

Fueled Weight 6391.6

Reference Area 200

Density 50000 ft 0.000364

0.45 0.009508 24229.9621 64.815703 552.5629


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Range and Endurance

65 of Endurance

Rangeof 24,230 miles

(21,055 )

Fuel mass fraction: 0.41

Globalhawk fuel mass fraction: 0.55


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Cruise Speed

Velocity at any Altitude given by Equation

Eq:

Know Density @ 50,000 ft = 3.6391 10

Filling in other known values Cruise speed = 372 mph

Cruise speed affects total endurance required

Higher cruise speed = less total endurance

Low cruise speed = more total endurance required


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Service Ceiling

Engine must be capable of producing enough thrust to fly

at 50,000 ft

Equation for maximum rate of climb

Eq: Service ceiling occurs where max R/C = 100

Want service ceiling to be higher than 50,000

Reduces time to climb from deployment altitude

Ensures aircraft can fly at desired altitude of 50,000 ft Service ceiling: 83,000 ft


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Landing Performance

We assumed touchdown velocity is 1.3

Calculated ground roll is 3460 ft

UAV can land on runways

Aircraft needs to be able to land on an aircraft carrier Aircraft carrier runway is typically 1000 ft

UAV can land on aircraft carrier with the use of a tail hook


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UAV Conclusion

Criteria/Constraints Globalhawk Predator Shadow 200 YRQ-0X

Endurance [hrs] 24 28 24 7.5 65

Service Ceiling [feet] 50,000 60,000 25,000 15,000 83,000

Range [nm] 24 hours + 3000 8,700 675 68 24,230Landing Distance [ft] 1000 for aircraft carrier 3460

10,000 for runway 3460

Cruise Speed [mph] 350 357 92 81 372


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Launch Vehicle Trajectory

Rocket will follow aelliptical, suborbital

trajectory between the

launch site and the target

Infinitely many ellipticaltrajectories, use trajectory

optimized for minimum

V, known as minimum

energy trajectory (MET)


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Launch Vehicle Trajectory

Spherical angle betweenlaunch site and targetlocation is the rangeangle

As range angle increases,the increasesexponentially

We chose to design a

rocket with range angleof 180


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Two Stage Analysis

Launch Vehicle First Stage IMF Second Stage IMFTitan II GLV 0.0525 0.0597Atlas F Centaur 0.0298 0.161Soyuz 11A510 0.0880 0.0992Delta 5920-8 0.0516 0.138

IMF Values found from 4 existing launch vehiclessimilar in payload and trajectory to our design.

Decided on IMF values of 0.0525 for the first stage,

and 0.0829 for the second stage.


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Split

Swept across thousandsof potential Voptions

to find the design with

the lowest launch mass.

Decided on a split of

57.63% provided by the

first stage

Percent provided by first stage

MassofLaunchV

ehicle(kg)


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Propellant

Kerosene and LOX propellant We chose to use kerosene for

its reliability and relative easeof storage when to comparedto fuels such as Hydrogen.

LOX has a relatively highvaporization point

Since LOX is cryogenic wedesigned our fuel tanks tohave an evacuated area

around the storage tank to actlike a thermos allowing foreasier storage.


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Propulsion System

The main criteria for the engines were a high ISP and

reasonable thrust using kerosene fuel.

For the first stage we decided to use the RD-191 engine

for the rocket. These engines had a specific impulse of 311 seconds and thrust of

2,079 kN at sea level.

This stage would require a propellant volume of 49.1 for

Kerosene and 89 for oxygen


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Propulsion System

Second stage has two RD-58M engines.

Chosen for high thrust and low mass

Use LOX and kerosene for fuel.

Each engine has a mass of 230 kg and thrust of 83.4 kN

Specific impulse is 353 seconds.

The fuel tanks for these engines would be a balloon

design that would support the tank structure by providing

outward pressure on the inside of the tank by some inertgas such as helium.


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Parachute Data

ROCKET UAVV IN M/S CHUTE DIAMETER [METERS] CHUTE DIAMETER [METERS]

1500 26.995 10.455

1450 27.457 10.634

1400 27.943 10.822

1350 28.455 11.021

1300 28.997 11.231

1250 29.572 11.453

1200 30.181 11.689

1150 30.831 11.941

1100 31.523 12.209

1050 32.265 12.496

1000 33.062 12.805

950 33.921 13.138

900 34.850 13.498

850 35.861 13.889

800 36.965 14.316

750 38.177 14.786

700 39.517 15.305

650 41.008 15.883

600 42.683 16.531

550 44.581 17.266

500 46.757 18.109

450 49.286 19.088

400 52.276 20.246

350 55.885 21.644

300 60.363 23.378

250 66.124 25.610

200 73.929 28.633

175 79.033 30.610

150 85.366 33.062

100 104.551 40.493

50 147.858 57.265


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Cost Analysis

$0

$2,000,000

$4,000,000

$6,000,000

$8,000,000

$10,000,000

$12,000,000

$14,000,000

$16,000,000

$18,000,000

$20,000,000

0 5 10 15 20 25

TotalCostofAllSys

tems

Number of Systems Purchased

Total Cost against Number of Systems Purchased

= 1 +1

Cost of One System:$12,179,460


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Cost Analysis

$0

$50,000

$100,000

$150,000

$200,000

$250,000

0 5 10 15 20 25

C

ostperHourofSurveillance

Number of Systems Purchased

Cost per Hour of Surveillance against Total Number ofSystems Purchased


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Launch Locations

Hickam AFB,Hawaii

Andersen AFB,Guam

Eglin AFB, Florida

Vandenberg AFB,California

Cape Canaveral,

Florida


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Conclusion

Done


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Image Sources

En.wikipedie.org/wiki/File:Fengyun-1C_debris.jpg

En.wikipedia.org/wiki/File:Debris-GEO1280.jpg

www.thespacereview.com/article/1323/1

www.geology.com/world/world-map.shtml

Documents

Design Proposal Presentation Final Slides