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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