1 HARP - High Altitude Reconnaissance Platform Design Proposal Dr. James D. Lang, Project Advisor Dr. Leland M. Nicolai, Project Sponsor Dr. Paul A. Wieselmann,

1

HARP - High Altitude HARP - High Altitude Reconnaissance Platform Reconnaissance Platform

Design ProposalDesign Proposal

Dr. James D. Lang, Project Advisor Dr. Leland M. Nicolai, Project Sponsor Dr.

Paul A. Wieselmann, Project Sponsor

Steven H. Christenson –Team LeadCeazar C. Javellana III Marcus A. Artates

2

PresentationPresentation OverviewOverview

-Define Requirements

-Design Process and Assumptions

-Aircraft Configuration/Sizing

-Weight Breakdown

-Mission Analysis and Compliance

-Aerodynamics

-Performance

-Propulsion

-Stability and Control

-Materials and Structure

-Cost Estimations

-Future Work

-References and Acknowledgements

3

RequirementsRequirements

Provide 24/7 ISR Coverage with 2 Aircraft

2000 nm Radius for ISR Mission

10500 nm Ferry Flight

6963 lb Payload (Installed Weight)

-(4) X Band Radar Arrays – 3.3 x 6.1 ft

-(2) UHF Radar Arrays – 4.9 x 40.6 ft

Minimize Take-off Weight and Life Cycle Cost

4

Mission Endurance

2*(One-Way Transit) + Time on Station

Time on Station

2*(One-Way Transit) + Turnaround Time

Derived Requirements for 24/7 Derived Requirements for 24/7 Coverage with 2 AircraftCoverage with 2 Aircraft

Transit Transit

TA

TOS

Transit

Transit Transit

Transit

TA

TOS

Transit

TOSTransit

TA

TOS Transit

Aircraft 1

Aircraft 2

Endurance

5

ISR MissionISR Mission

Descend to Sea Level

Climb to Cruise

Altitude

Cruise Out 2000 nm Cruise Back 2000 nm

Loiter 16 Hours (TOS)

Sea Level Loiterfor 30 min

55000 ft

Distance (nm)

2000 nm

6

Max Distance Ferry Mission Max Distance Ferry Mission

Descend toSea Level

Climb toCruise

Altitude

Cruise 10500 nm

Sea Level Loiterfor 30 min

55000 ft

Distance (nm)10500 nm

7

Assume Wto and

W/S

Size Wing

Calculate Component Weights

Calculate Fuel Fractions

Yes/NoDetermine Fuel

Available

Fuel)aval

> Fuel)reqd

Determine Fuel Required for

Mission

Aerodynamics Size Engine Performance

AR, Taper, Sweep

Fuselage Sizing and Shape

Estimate Tail Size

Study Mission Requirements

Refine Wto and

W/S Estimates

Refine Aerodynamic Parameters

Size Control Surfaces/Tail

Calculate Drag

Determine Performance Capabilities

Mission Requirements

Met?

Refine Wto and W/S

Optimize Design

-Assumptions Made/Refined-

-Configuration Assumptions Made/Refined to Meet Mission Requirements-

Design ProcessDesign Process

Yes/No

8

Aircraft ConfigurationAircraft Configuration

-L/D)max,wing = 35 for 0 deg Sweep, 20 AR, 60% Laminar Flow

Lockheed Martin Aerodynamic Data

-2250 lb Thrust, .55 TSFC for 2015 Advanced Technology Turbofan Engine at Full Power and 55000 ft

Design Analysis Based on the Following Assumptions:

9


Wto = 50000 lb W/S = 60 lb/ft^2

Wing Area = 833 ft^2 Wing Span = 129 ft

Wing Sweep = 0 deg Aspect Ratio = 20

10

Radar GeometryRadar Geometry

X Band Radar (4)

-3.3 x 6.1 ft

-Azimuth Field of Regard (FOR) +/- 70 degrees

-Located to give 360 Degree Coverage

UHF Radar (2)

-4.9 x 40.6 ft

-Azimuth FOR +/- 70 degrees

-Located to View Out Each Side

11

Horizon DistanceHorizon Distance

55000 ftHorizon

5.17 deg250 nm LOS

Design Array Angles for Desired Footprint

12


Wing Area = 833 ft^2 Wing Span = 129 ft

Wing Sweep = 0 deg Aspect Ratio = 20

Fuselage

Length = 62 ft

Height = 6 ft

Width = 10 ft

13


14


Wing Fuel Tank

Center of Gravity

& Aerodynamic

Center

15

1. Start up/Take-Off .970

2. Climb to Cruise Alt .950

3. Cruise Out .902

4. Loiter on Station .754

Loiter Fuel 10219 lb

Maneuvering Fuel 671 lb

5. Cruise Back .902

6. Descend to SL 1.00

7. Loiter 20 min .994

Take-Off Weight 50000 lb

Fuel Weight 23874 lb

Fuel Fraction .48

Fuel Volume 3511 gal

Weight Fractions -ISRWeight Fractions -ISR

-Cruise at .943*L/D)max

-Loiter at L/D)max

(1) 2015 Technology Turbofan Engine

SLS Thrust = 8000 lb

SLS TSFC = .40

T/W = .16

16

-16 Hour TOS-

Cl = .864 L/D)max = 31.52

Mach .6 and 55000 ft

ISR Mission ComplianceISR Mission Compliance

Mission Endurance

2*(One-Way Transit) + Time on Station

= 2*(5.52) + 16.2 hr = 28.4 hr

Time on Station

2*(One-Way Transit) + Turnaround Time

= 12.2 hr + 4 hr = 16.2 hr

-Two Aircraft Coverage-

-2000 nm Range-

Cl = .628 L/D = 29.72

Mach .6 and 55000 ft

Total Mission Fuel Required: 23874 lb = 3511 gal

17

Weight Fractions - FerryWeight Fractions - Ferry

1. Start up/Take-Off .970

2. Climb to Cruise Alt .950

3. Cruise 10500 nm .567

5. Descend to SL 1.00

6. Loiter 20 min .994


Fuel Weight 24685 lb

Fuel Fraction .49


Design Pushed by 10500 nm Ferry Flight

Approx 800 lb Additional Fuel Required

18

AerodynamicsAerodynamics

Aspect Ratio = 20 Span = 129 ft

Wing Sweep = 0 deg e = .9

t/c = .15 K = .01768

Taper Ratio = .50 MAC = 6.7 ft

Croot = 8.6 ft Ctip = 4.3 ft

Airfoil: Modified Lockheed Martin Sensorcraft Wing15% to Provide 60% Laminar Flow

19


L/D)max,wing = 35 Lockheed Martin Aerodynamics Data

Cdo)wing = .00817 Referenced to Sref

Cdo)fuselage = .00369 Referenced to Sref

Cdo)tail = .00121 Referenced to Sref

Cdo)aircraft = .01393 Calculated with Interference Effects

L/D)max,aircraft = 31.52 From L/D vs Cl Plot

20


Cl = .864 for L/D)max and Minimum Drag

Clalpha = 6.9 rad-1 = .12 deg-1 at Mach .6

Stall Velocity Based on Cl)max of 2.0

Candidate High Lift Devices

-Mission Adaptive Wing (MAW)

-Trailing Edge Flaps

21


Fuselage Sized to Hold Radar Arrays

Length = 62 ft

Depth = 6 ft

Width = 10 ft

Fineness Ratio = 6.2

Volume = 2922 ft^3

Wetted Area = 1067 ft^2

Max Cross Sectional

Area = 47 ft^2

22


Cl/Cd vs Cl

0.00

4.00

8.00

12.00

16.00

20.00

24.00

28.00

32.00

0.00 0.25 0.50 0.75 1.00 1.25 1.50 1.75 2.00 2.25 2.50

Cl

Cl /

Cd

Cl / Cd

Cd = Cdmin + K''(Cl - Clmin)^2 + K'Cl^2Cdmin ~ Cdo = .01393

Clmin =.4K'' ~ .00122, K' = .0177

L/D)max = 31.52

23


Drag Polar at Mach .6

0.00.10.20.30.40.50.60.70.80.91.01.11.2

0.000 0.005 0.010 0.015 0.020 0.025 0.030 0.035 0.040 0.045 0.050

Cd

Cl

Drag Polar

Calculated Cl = .864 for L/D)max

24


-Insufficient Data in References to Accurately Calculate MDD

-Concern that at Cruise Velocity and Altitude (M .6 @ 55000 ft) Airfoil is Near MDD

-Supercritical Wing

MDD, Drag Divergent Mach Number

25

PerformancePerformanceLimit Load Factor 1.25

Ultimate Load Factor 1.88

Turn Load Factor 1.15

Maneuvering Turn Rate 1.8 deg/s

Dynamic Pres Limit 450 lb/ft^2

Stall Velocity 159 ft/s

Take-Off Velocity 191 ft/s

Take-Off Distance 5000 ft

Landing Distance 4000 ft

Braking Acceleration –7 ft/s^2

26

PerformancePerformance

27


28


29

PropulsionPropulsion2015 Technology Turbofan Engine

Moderate Bypass Ratio

8000 lb Thrust (Sea Level Static)

.40 TSFC (Sea Level Static) Dimensions:

Length 115 in (9.6 ft)

Diameter 41 in (3.4 ft)

Engine Weight: 1600 lb

System Weight: 3100 lb

-Pitot Inlet, 10 ft^2 Capture Area

-Fixed Convergent Nozzle, 6 ft^2 Exit Area

30

PropulsionPropulsion

Full Throttle Engine Thrust

0

2,000

4,000

6,000

8,000

10,000

12,000

14,000

16,000

18,000

20,000

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

Mach Number

Thru

st [l

b]

SL

5k

10k

15k

20k

25k

30k

35k

40k

45k

50k

55k

60k

31


Full Throttle Engine SFC

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

Mach Number

SFC

[lbm

/hr/

lb]

SL

5k

10k

15k

20k

25k

30k

35k

40k

45k

50k

55k

60k

32

Thrust Data vs Altitude at Mach .6

1000

3000

5000

7000

9000

11000

13000

0 10000 20000 30000 40000 50000 60000

Altitude [ft]

Th

rust

[lb

f]

Thrust at Mach .6


33

Auxiliary PowerAuxiliary Power

Required Power 128 kW

Power Available from Engine 70 kW = .061*Talt

Additional Power Required 58 kW

Total Weight 1304 lb

APU Fuel Weight 595 lb

Total Weight 1899 lb

APU – Continental L/TSIO-360

34

Auxiliary PowerAuxiliary Power

Engine Excess Power

kW = .061*Talt

Additional Thrust 957 lb

Additional Fuel 8562 lb

(T-D)*V = Power

Additional Thrust 58 lb

Additional Fuel 523 lb

Average Additional Fuel 4542 lb

35

Fuselage 3415 lb

Wing 4928 lb

Control Surface(s) 2508 lb

Tail 297 lb

Landing Gear 1677 lb

Propulsion System 3100 lb

Flight Systems 460 lb

Fuel System/Tanks 496 lb

Hydraulic System 172 lb

Electrical System 849 lb

Air Cond/Anti-ice Sys 794 lb

Payload (Installed) 6963 lb


Empty Weight 18697 lb

Weight with Payload 25660 lb

Fuel Weight Available 24340 lb

Fuel Fraction .49


Weight Build-upWeight Build-up

-Fuselage and Landing Gear Weight Reduced by

15% and 5%, respectively, for 2015

Technology Target Factors

36

Stability and ControlStability and Control

Center of Gravity and Fuel Schedule

Fuel Consumption Schedule

% Fuel Fuel 1 (Wing) Fuel 2 (Fwd) Fuel 3 (Aft)100 100% 100% 100%80 80% 80% 80%60 65% 60% 55%40 60% 40% 30%20 40% 20% 10%5 5% 5% 5%

Center of Gravity Wto 50000

100% Fuel

Component Weight [lb] Dist [ft] Moment [ft-lb]Fuselage and sys 7366.47 31.00 228360.61

Fuel 1 (Wing) 8278.49 31.35 259544.36Fuel 2 (Fwd) 8278.49 19.00 157291.30Fuel 3 (Aft) 8278.49 43.00 355975.04

Payload 6963.00 28.00 194964.00Wing and Cont Surf's 7436.49 29.20 217145.57

Dist to Wing C/4 7436.49 31.35 233146.35Horiz Tail 214.29 57.00 12214.35Vert Tail 83.79 57.00 4775.91

Engine 1 3084.00 43.50 134154.00

Sum Wt Sum Mom49983.51 1580425.91

Distance from Nose Xcg [ft] 31.62

37


Static Margin (SM) Summary

Center of Gravity Travel

0

20

40

60

80

100

31.4 31.5 31.6 31.7 31.8 31.9 32.0

Xcg [ft]

% F

uel

CG Travel Aerodynamic Center

Fuel Capacity 100% 80% 60% 40% 20% 5%Xac - Xcg [ft] 0.2109 0.1567 0.2069 0.2755 0.2076 0.3082

Static Margin, SM 0.031 0.023 0.031 0.041 0.031 0.046

Average SM 0.034 Cmalpha -0.23

38


Moment Coeficient vs Angle of Attack

-10

-5

0

5

10

-5 -3 -1 1 3 5 7

AoA[deg]

Cm

δe=0 δe=+3 δe=+1 δe=-1 δe=-3

Cmo = .0681

39


Ailerons

Area = 37.9 ft^2 each

MAC = 1.47 ft

Span = 25.8 ft

Flap Chord: 25% Wing Chord at Root

Flap Span: 27% of Wing Span

Flaps

Area = 38.0 ft^2 each

MAC = 2.15 ft

Span = 17.7 ft

Total Control Surface Area: 152 ft^2

Aileron Chord: 22% of Wing MAC

Aileron Span: 40% of Wing Span

40


V-Tail

Cvt = .0145 Svt = 55.7 ft^2

Cht = .34 Sht = 67.7 ft^2

42 deg from Vertical

Rudder Area = 18.6 ft^2 = (1/3)Svt

41

Materials and StructureMaterials and Structure

Carbon Fiber -Wings-Control Surfaces-Fuselage

Fiberglass-Array Panels

Material Selection

Structural Concept

Semi-Monocoque Fuselage Structure

Carbon Fiber Wing Box, Spars and Landing Gear Struts

42


Airload Distribution

0

20

40

60

80

100

120

0 100 200 300 400 500 600 700 800

y [in]

Air

load

Inte

nsi

ty [l

b/in

]

43


Moment Distribution

-1.8E+07-1.6E+07-1.4E+07-1.2E+07-1.0E+07-8.0E+06-6.0E+06-4.0E+06-2.0E+060.0E+00

0 200 400 600 800

y [in]

Mo

me

nt

[lb

/in]

44


Shear Distribution

0

10000

20000

30000

40000

50000

60000

0 200 400 600 800

y [in]

Sh

ea

r F

orc

e [

lb]

45


Ixx = 2.89E3 slug-ft^2

Iyy = 1.93E5 slug-ft^2

Izz = 6.86E5 slug-ft^2

Mass Moments of Inertia Based on Historical Data

46

Cost EstimationsCost EstimationsEngineering Hours, Tooling Hours, Manufacturing Hours and Manufacturing Material Costs Based on Historical Data and: -Number of Aircraft Produced -Aircraft Take-off Gross Weight -Maximum Velocity

Flight Test Costs Based on Historical Data and: -Number of Flight Test Aircraft -Aircraft Take-off Gross Weight -Maximum Velocity

Quality Control Hours Based on Historical Data and: -Manufacturing Hours

Development Support Cost Based on Historical Data and: -Aircraft Take-off Gross Weight -Maximum Velocity

Engine and Avionics Cost Provided By: -Lockheed Martin

47

Cost EstimationsCost Estimations

Hours

Engineering 7,568,054

Tooling 4,483,622

Manufacturing 13,472,465

Quality Control 1,791,838

Aircraft to be Procured: 100

Flight Test Aircraft: 6

Costs

Development Support 88,831,854

Flight Test 57,056,356

Manufacturing Materials 260,106,607

Engine 206,700,000

Avionics 1,590,000,000

Labor Rates Adjusted to 1999 Dollars

Engineering $85

Tooling $88

Manufacturing $73

Quality Control $81

Estimated RDT&E + Flyaway Cost = $4,470,179,979

44. 7 Million / Aircraft

48

Future StudyFuture Study

-Tailor Fuselage Shape to Minimize Flow Separation

-Analyze Control and High Lift Concepts Mission Adaptive Wing (MAW)

-Analyze Desired Radar Footprint for Exact Array Orientation

-Wing Dihedral

-Low Observables

-Possible Requirement for Satellite Antenna

System Configuration

49

Future StudyFuture Study

-Utilize VaRTM Technology

-Incorporate High Strength Composites to Replace Traditional Metal Components

-Refine Installed Thrust Data

-Refine Inlet/Nozzle Design

Performance

Cost

50

References and References and AcknowledgementsAcknowledgements

References:

Fundamentals of Aircraft Design, Nicolai, L.M., Revised 1984

Lockheed Martin Aerodynamic Data, Nicolai, L.M.

Aircraft Design: A Conceptual Approach, Raymer, D.P., Third Edition

Acknowledgements:Acknowledgements:

Dr. James D. Lang, Project AdvisorDr. James D. Lang, Project Advisor

Dr. Leland M. Nicolai, Project SponsorDr. Leland M. Nicolai, Project Sponsor

Dr. Paul A. Wieselmann, Project Sponsor

51

Thank YouThank You

Documents

1 HARP - High Altitude Reconnaissance Platform Design Proposal Dr. James D. Lang, Project Advisor Dr. Leland M. Nicolai, Project Sponsor Dr. Paul A. Wieselmann,