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Design and Analysis of the Jayhawk Economic Turboprop

Transport (J.E.T.T)

AIAA Undergraduate Team Aircraft Design Competition

2013 - 2014

Brandon Basgall

Katie Constant

Eleazar Lachino

Adrian Lee

Emily Thompson

Alejandra Escalera

Instructor: Dr. Ron Barrett

Department of Aerospace Engineering

May 10, 2014

a541l091

Sticky Note

This is an Except from University of Kansas Undergraduate Team-Alpha Aircraft Design Report submitted for the 2013-2014 American Institution of Aeronautics and Astronautics (AIAA) Undergraduate Team Aircraft Competition, in which the University of Kansas Team-Alpha won first overall placing among 18 other teams from different campus around the globe. Note this is an except not the actual full report.

Aerospace Engineering Department 3

Mission Specification and Profile

Provided by AIAA during the fall of 2013, the following outlines the RFP guidelines to be adhered to in the

design of a new and upcoming turboprop regional airliner. The following lists the major characteristics needed to be

considered for the newly designed regional turboprop aircraft:

Table 2-1: Regional Turboprop Mission Specifications and Requirements (Ref. 2)

Economic Mission Requirements

Available Passenger Seating 75

Range 400 nautical miles

Design Mission Requirements

Available Passenger Seating 67

Range 1600 nautical miles

Additional Considerations Necessary Mission Fuel Reserves

Performance Requirements

Take-off Runway Length at MTOW 4,500 feet (or shorter)

Take-off Conditions (Altitude and Temp.) 7,800 feet at 85° F

Cruise Speed (Mach) 0.62 – 0.68

Cruise Altitude 25,000 – 31,000 feet

Fuel Consumption 65% fuel reduction from similar currently operational

regional jets

Cost

Production Run 15 years period, 400 aircraft

Total Cost Must be substantially less than currently operational

regional jets

Entry into Service 2022, minimum lifetime of 30 years

Additional Considerations

Passenger Comfort and Acceptance “Jet-like” Experience

Passenger with Baggage Weight 225 lbs


From the previously

defined variables to be included

in the design process of the new

regional turboprop aircraft, the

following mission profiles have

been created. Figure 2-1 shows

the defined economic mission

under normal runway

considerations of 4,500 foot

runways. From this mission

profile, it can be seen that the

cruise conditions for this profile are noted as cruise speed between Mach 0.62 and Mach 0.68 with a cruising altitude

between 25,000 and 30,000 feet.

To accommodate the requested design mission of 1,600 nautical miles carrying 67 passengers, Figure 2-1 also

displays the mission profile

for the conventional design

mission. Within this mission,

the cruise speed and altitude

will remain the same as seen

with the economic mission, as

well the same 4,500 foot

runways. With the design

range in mind, the design of

fuel reserves will be taken

into account. As designated within the RFP, the design fuel reserves will meet the FAR 25 Loiter regulations, stating

the full 45 minute loiter fuel-reserve capacity.

To be able to achieve the designated “high-hot” conditions outlined in the AIAA RFP, Figure 2-2 displays the

adjusted mission profile. To adjust for the “high-hot” conditions, the runway length will be adjusted to 8,000 feet

Figure 2-1: Design and Economic Mission Profiles of Turboprop

Aircraft

Figure 2-2: Design and Economic Mission Profiles for "High-Hot"

Conditions


Figure 4-3: Determination of Design Fix Date with Respect to

Variables (Ref. 1)

equation to model the distribution has been

determined, it will be placed on the same plot

as the first flight distribution.

The vector analysis process gives

engineers a tool with which they can track

and predict any vector of engineering design

variables at any given point in time. This

point can be either in the past to analyze

previous trends or in the future to see what

they should prepare for. The process makes

use of distribution curves by taking into

account all of the desired design variables for

as many aircraft, of the similar class as the

one being asked for. The following is the

procedure that was performed by the contestants to establish some of the critical design variables (Ref. 21 )

First a product timeline must be developed using aircraft of the same class as the one being designed. The

most important dates of an aircraft’s lifecycle must be researched and gathered for use. The date of first flight of the

aircraft will be plotted onto a graph and

a distribution curve will be generated to

fit the data. The peak of this, also the

highest year for first flights of aircraft

will be labeled as year zero of the

aircraft’s lifecycle. The res t of the

distribution will be positive or negative

years with respect to the reference zero.

Once this distribution has been

established the design fix date for the same aircraft, it will generate another distribution curve. Once again the year

at which the distribution reaches its peak will be labeled as zero. Once the graph has been observed and a good

Figure 4-2: Determination of Statistical Interval

Related to Rvi (Ref. 21)


Empty-to-Takeoff Weight Ratio, We/Wto (~)

Mar

ket

Shar

e, (

%)

1983 1984

1987

1988 1989

1990

1991

1985 1986

1992 1993

1994 1995

1996 1997

1999 1998

2000 2001

2002 2003

2004 2005

2006

2007 2008

2009 2010

2011 2012

2013

2020

0.000

0.100

0.200

0.300

0.400

0.500

0.600

0.700

0.800

0.5 0.6 0.7

Pe

rce

nta

ge o

f M

arke

t Sh

are

, (%

)

Empty-to-Takepff Weight Ratio We/Wto (~)

ATR-42 33%

Fairchild 37%

EMB-120 23%

F-27 5%

1986 Weight Standard Deviation

BAE ATP 2%

Figure 5-1: STAMPED Weight Analysis


6.2 Sizing Chart Analysis

Utilizing the STAMPED procedure for identifying a

preliminary design point, the following table is a summary of the

first iteration of design. Additional analysis and multiple iterations

through Class I and Class II design would ultimately lead to the

final design sizing selections. Figure 6-1 is a visual representation

on how the sizing values were determined. As the figure shows, by

documenting the changes through time, the design point was determined.

Figure 6-1: Statistical Time and Market Predictive Engineering Design (STAMPED) Sizing Chart Results

In addition to the STAMPED analysis, a sizing chart was constructed following the procedure outlined in

Reference 3. The results from these calculations are presented in the figure below. The results from the two methods

Area, S (ft2) 800

Power (shp) 7530

Wing Loading (psf) 80

CLmax,clean 1.9

CLmax,landing 2.1

CLmax,TO 2.6

Table 6-1: Sizing Chart Results


Take off Wing Loading (W/S)TO (lbf/ft2)



1998 Example

σ = +1

σ = -1

σ = -1

σ = +1

2020

Standard

Deviation at σ

= ±1

Design

Point

W/S

Sizing Chart

W/P

W/S

Po

wer

Lo

adin

g,

(W/P

) TO (

lbf/

hp

)



1998 Example

Po

wer

Lo

adin

g,

(W/P

) TO (

lbf/

hp

)

Take off Wing Loading (W/S)TO (lbf/ft2) Po

wer

Lo

adin

g,

(W/P

) TO (

lbf/

hp

)

W/P

Figure 6-3: STAMPED Power and Wing Loading


Class I Sizing

The following sections will develop the calculations completed for a class I design analysis. This section will

begin with a range of proposed configurations, followed by a down-selection to four final preliminary configuration

designs. Class I design was completed on each of these four designs the results of which determined the design best

fit for Class II analysis.

7.1 Preliminary Configuration Selection

7.1.1 Selection of the Aircraft Configuration

For this initial stage in aircraft configuration, all possible options were considered. Driving factors of

configuration selection included technical considerations such as C.G. excursion, and fuel capacity but also design

goals such as range, jet-like experience for passengers, and low procurement cost.

7.1.2 Aircraft Configurations and Merit Criteria

The following is a list of 12 different aircraft configurations that were considered by the team in the

preliminary configuration selection process. These various designs were evaluated with respect to the requirements

and objectives outlined in the RFP (Ref. 2).

Table 7-1: Air Configuration Considerations

1. Box Wing

2. Vertical/Short Take Off

& Landing (V/STOL)

3. Delta Wing

4. Conventional (High-

Wing, Conventional Tail

5. Tail Boom, Swept Wing 6. Three Surface

7. Flying Wing 8. Tandem Wing

9. Inverted Gull Wing

10. W Wing

11. Channel Wing

12. Double Fuselage


Figure 8-8: Isometric View of Engine Access Panel

propeller in respect to the wing was chosen specifically for geometric clearance during low speed approach with

bank angles around 5° as well as

during take-off rotation (Ref. 4).

The shroud will have 5 struts for support and

withstand the vibration generated by the propeller and engine.

Figure 8-9: Engine Removal Method

The engine removal from the shroud and nacelle will require a special pulley as shown in Error! Reference source

not found.

Figure 8-7: Side View of Shrouded Engine


on the ground. The following figures demonstrate the Prandtl-tailored δ3 wingtips which the team opted to

implement. Prandtl with his research found in Über Tragflügel kleinsten induzierten Winderstandes, that the wing

weight could be reduced by minimizing the moments, either creating negative bending moments to counteract the

positive moments or forcing them to approach zero (Ref. 24). Figure 8-12, from Prandtl’s Über Tragflügel kleinsten

induzierten Winderstandes shows the span loading concepts that were established with his research. Prandtl made

the following conclusions in his document: elliptical distribution of lift is disadvantageous if trying to minimize

induced drag, although the exact shape of sharp wingtips is not of high importance, sharp wingtips are advantageous

when compared to squared wingtips (Ref. 24).

Keeping in mind the conclusions drawn by Prandtl’s studies, the target wing span with wing tips folded

could grow no larger than 79 feet (Category 2 regulations). For optimal L/D, a total wing span was set at 118 feet

with an aspect ratio of 23.

Figure 8-13: Boeing 737 Clean Configuration, With Prandtl-tailored δ3 Wingtips both Extended and Folded

and Folding Mechanism (Ref. 24)

Clearly, 118 foot wing span and aspect ratio is far outside the allotted gate spacing thus need for foldable wingtips

becomes necessary to make this possible. One of the reasons why Prandtl’s concept has not been implemented in

current aircraft is the weight of the hinge and actuator to fold the wingtips (Ref. 25). While the overall performance

is increased, the added weight makes this design impractical for use. However, with the implement

ation of Pressure Adaptive Honeycomb (PAH) weight concerns are eliminated. In this application, Pressure

Adaptive Honeycomb would be used as the actuator of the wingtip and for the implementation of a δ3 hinge

ensuring that the moment at the wing tip is equal to zero or insignificant (Ref. 25). While an advanced technology,

PAH can be produced at a low manufacturing cost from aluminum and nylon pouches and can be integrated to

conventional structures in the aerospace industry while maintaining a low part count (Ref. 24). Although the

Pressure Adaptive Honeycomb can be activated by the external ambient pressure, it can be designed to actuate the


Figure 9-9: Carpet Plot for Wing Loading at a Specific Range

Table 9-7 : Carpet Plot Design Parameters

The carpet plot here is serving as a verification of the

design parameters selected. As Figure 9-9 shows, the

parameters selected for the design of the J.E.T.T. can be

verified and are at the optimal design point for the RFP. The

following table is a summary of the design parameters

collected from the above carpet plot.

0.298

0.3

0.302

0.304

0.306

0.308

0.31

74 76 78 80 82 84 86 88

Spe

cifi

c R

ange

, SR

(n

mi/

lb)

Wing Loading, W/S (lbf/ft^2)

AR = 25

AR = 23AR = 21

AR = 19

Wf = 5300 lbf

Wf = 5400 lbf

Wf = 5500 lbf

Wf = 5600 lbf

Wf = 5700 lbf

Design Point

Insufficient Fuel to reach 1600nmi

Insufficient Fuel Space on Wing

Parameter Value

Specific Range, SR (nmi/lb) 0.305

Weight of Fuel, Wf (lb) 5500

Aspect Ratio, AR 23

Lift-to-Drag Ratio, L/D 21.1

Wing Loading, W/S (lbf/ft^2) 80


Figure 10-1: All Systems and Subsystems of the J.E.T.T

Layout of Systems and Subsystems

All of the major systems were modeled in this aircraft which included:

Flight Control System

Hydraulic System

Water and Waste System

Environmental System

Electrical System

Fuel System

De-Icing System

10.1 Flight Control System

The J.E.T.T. is equipped with an irreversible flight-by-light control system. An irreversible system was

chosen for obvious reasons and an irreversible flight control system in an aircraft of this size would be extremely

impractical and outdated. Flight-by-light was chosen in place of the more traditional fly-by-wire since the

technology is available now and will only be more advanced in years to come. As it stands currently, the optics in

fly-by-light systems require less space and in an aircraft as constrained for extra space as the J.E.T.T. this is

important. Additionally, the 18 fiber optical cables are designed to meet all environmental, mechanical and optical

requirements set for by FAR requirements (Ref. 30). For safety purposes, this system is quadruple


10.6 Water and Waste System

The water and waste system will be housed in the belly of the aircraft along with many of the other systems.

Waste that is produced in either the front or the rear of the aircraft will be pumped into either one of the holding

tanks which upon landing will be emptied. Since the entire fuselage will be pressured and heated, there is no concern

that water lines may freeze in this design.

10.7 Environmental System

The final system in this aircraft is the environmental system which includes the air conditioning and oxygen

system; both extremely important for day to day operations and are both doubly redundant. As was mentioned

previously, both the APU and RAT are capable of supplying power to the oxygen system if failure occurs elsewhere.

Figure 10-8: Environmental System Shown in Green Provided Oxygen and Air Conditioning

The J.E.T.T. will be equipped with an open-loop redundant cabin air conditioning and pressurization system.

For passenger comfort, the cabin will be pressured at 7,500 feet (11.1 psi) and temperatures will be maintained

around 70 degrees Fahrenheit.


Inboard Profiles

From the required RFP of a maximum amount of passengers to carry, the J.E.T.T. can accommodate 76

passengers with the addition of 2 flight attendants. To get this arrangement, the seating is organized with a 4-abreast

seating throughout the full cabin. This seating arrangement allows for space at the front and the aft of the cabin for a

fully functioning galley and a lavatory at the tail of the aircraft. Also, the symmetric arrangement of the seats fed

into the considerations of carry-on baggage volume, aisle width and length, as well as the ease of redesigning for

family variants in production. The following table details the specific cabin characteristics.

Table 11-1: Internal Cabin Characteristics

Inner Fuselage Diameter 102 in Number of Seats 76

Seat Pitch 20.0 in Crew Seats Attached to Aft Bulkhead Wall

Seat Width 19.5 in Galley Floor Size 39 in x 26 in

Aisle Width 19.0 in Lavatory Floor Size 39 in x 47 in

Maximum Aisle Height 75.0 in Carry-On Baggage Volume 100.7 ft3

Aisle Width between

Overhead Compartments

23.65 in Tail Cone Baggage Volume 234.7 ft3

To adhere to FAR 25 regulations, a Type III sized door will be placed in the rear of the cabin for emergency

exits. The other two doors at the front of the fuselage are classified as Type IV doors to aid in loading and loading of

passengers. The duel door system also allows the aircraft to be unloaded on either side of the fuselage, allowing

versatility for airport handlers. Lastly, a baggage door is accounted for in the fuselage cone, where a majority of the

baggage will be

stowed during

flight. In the

following figure,

the J.E.T.T.’s

internal cabin is

displayed with the

fully arranged

components.

Cockpit

Galley

Overhead

Baggage

Compartments

Lavatory

Tail Cone

Baggage

Door

Emergency

Egress Door

Figure 11-1: Internal Layout of Major Components of J.E.T.T.


11.1 Accessibility

The J.E.T.T. is designed for rapid “turn-around”, accessibility, maintainability and inspectability. The aircraft

was laid out such that it would be possible for multiple operations to occur at once such as the loading and unloading

of passengers, refueling, replenishing of potable water, cleaning of the airplane cabin, servicing lavatories as well as

others. The following Figure 11-2 presents a diagram that demonstrates the potential of the airplane to satisfy

customer’s needs and requirements.

Figure 11-2: Terminal Servicing of J.E.T.T.


Structures, Manufacturing, & Production

12.1 Structural Design

Integration of the aircraft structure is an integral aspect of design and must completed carefully to avoid the

addition of weight that is unnecessary and serves no added benefit to performance both in flight and in structural

stability.

12.1.1 Wing Structural Design

Similar to most designs, two main spars (wing box) run along the length of the wing and all sections not

including the folded winglets. The winglets are supported with pressure adaptive honeycomb and require no

additional internal structure. The front spar is located at 18% of the chord while the rear spar is at 72% of the chord.

Wing ribs were placed where the structure required the most support and then every 19 inches. The fuel tanks were

then integrated between the rib structures accordingly. Wing stiffeners were placed based upon recommendations

provided in Reference 8 on general arrangement if he wing and stiffness of the aluminum skin.

For the wing – fuselage integration, the wings

are bolted to a fuselage carry through section

near the top of the fuselage.

While the wings on this aircraft are

large, with an aspect ratio of 23, wing weight

is not excessive therefore all internal structure

as well as skins will be manufactured from

composite materials. The following images depict the wing structure and design.

12.1.2 Empennage Structural Design

Much like the wing, the empennage spar locations

are at 18% and 72% of the chord. Rib spacing was

selected to be 20 inches. Stringers again can vary and were

placed appropriately to best support the structure. Also

similar to the wing and which will be found throughout,

the skin will be composite materials.

Figure 12-1: Wing Structure and Fuselage Integration

Figure 12-2: Empennage Structure and Fuselage

Integration


Figure 12-4: J.E.T.T. Complete Aircraft Three View


12.1.4 Proposed Manufacturing

Similar to current industry, the J.E.T.T.

will be assembled in several outsourced locations

around the world, and finally assembled in a

domestic facility. From current processes used by

Boeing and Spirit Aerosystems, the J.E.T.T. has

several major sections of the aircraft that are

composed of composite materials, thus regulated

facilities that specialize in these processes would

be preferred. Table 12-1 displays the J.E.T.T. broken down into the major sections defined with the main material

defined, as well as the location in which the component will be produced

Table 12-1: Designation of Major Components with Material Selection and Location of Production

Designation Component

Name

Main Material Production

Location

1 Cockpit Composite Outer Skin/

Aluminum Main Structure

United States

2 Cabin Doors Aluminum Main Structure France

3 Fuselage Composite Outer Skin/


Japan

4 Engine Nacelles Boron, Kevlar Canada

5 Wing Composite Outer Skin/


Japan

6 Control Surfaces Composite Outer Skin/


Australia

7 Flexing Winglets Composite Outer Skin,

Honeycomb Main Structure

United States

8 Fuselage Cone Composite Outer Skin/


United States

9 Vertical Tail Composite Outer Skin/


United States

10 Horizontal Tail Composite Outer Skin/


Italy

11 Inside Seating Fire Retardant Materials France

12 In Flight Entertainment Electrical Components Japan

13 Cockpit Avionics Electrical Components United States

In addition to displaying what main material has been selected with the location of production, Figure 12-5

displays the attachment schematic of the full J.E.T.T. once all components are received at a domestic location.

1 2

3 4

5

6

7

8

9

10

Figure 12-5: Designation of Main Components of J.E.T.T. for

Production


Wing Structure Middle Fuselage

Structure

Aft Fuselage

Structure

Front Fuselage

Structure

Vertical Tail

Structure

Horizontal Tail

Structure

Assembly of Main

Aircraft Structure

Engine Mounting

System Integration

Internal Cabin and

Seating Integration

Main/Nose Landing Gear

Integration

Final Assembly with

Main Skin Components

and Flexing Wing Tips Complete Assembly of

J.E.T.T.

Figure 12-6: Manufacturing Schematic


Advanced Technologies

Several advanced technologies were employed in this design to meet performance and mission objectives. A

summary of each of the technologies as well as how and why they have been implemented is discussed below.

14.1 Pressure Adaptive Honeycomb (PAH)

The first of the advanced technologies has been a developing technology for many years but has yet to be

actively implemented. Pressure Adaptive Honeycomb (PAH) is an effective approach to minimizing structure

weight but continuing to perform to the same caliber. The following section explains how PAH functions and what

makes it so effective.

CDP is the difference between cell pressure and ambient pressure, in which the cell pressure can be controlled

by a pressure source within the aircraft. The preferred source of pressure is to have a connection between the

Pressure Adaptive Honeycomb structure and the compressor stage of the engine; the PAH will not negatively affect

the performance of the engine as it does not require a continuous flow of air (Ref. 27). The redundancy in the system

comes from the use of CO2 cartridges if the engines fail during flight (Ref. 27).

Figure 14-1: Honeycomb Grid Breakdown: Structure-Cellular Tissue-Single Cell-Cell Wall (Ref. 27)

Figure 14-2: Pressure Adaptive Honeycomb Actuation and Structure Possibilities (Ref. 27)

The ideal Pressure Adaptive Honeycomb would have the following characteristics as described in reference

25: easily controlled, capable of handling (-50+%) strains, stiff and strong enough to handle “real” loads, FAR

23/25 certified, all substructure within the fatigue zone labeled as “infinite life”, no continuous/dedicated power


Table 14-1: Research Development, Testing and

Evaluation Costs

Cost Estimation and FlyOffs

14.4 Cost Estimation

Following the completion of the J.E.T.T design, a detailed cost analysis was completed to estimate and

predict not only the acquisition cost of this aircraft but also expenses associated with research, testing and

development as well as operating

costs. This analysis was completed

using calculation methods as

outlined in reference 10 which

emphasizes the different costs of

an aircraft as it relates to the

different periods along an aircraft

program Life Cycle as shown in

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self-reference.. As established in the RFP, when this design enters production it should be expected that a minimum

of 400 units will be manufactured. Additionally, operational costs are largely dependent upon mission length;

therefore costs will differ between the economic versus design missions.

14.4.1 Research, Development, Testing and Evaluation Costs (CRDTE)

The first phases of the life cycle of an aircraft include activities which develop a new aircraft design from

first concept, planning and concept exploration to certification. This phase accounts for a substantial amount of the

life cycle costs and is broken down into the following

sections:

Airframe Engineering and Design Cost, Caedr

Development Support and Tsting Cost, Cdstr Flight Test Airplanes Cost, Cftar Flight Test Operations Cost, Cftor Test and Simulation Facilities Cost, Ctsfr RDTE Profit, Cpror Finance Costs, Cfinr

Caed_r $ 186,367,204.51

Cdst_r $ 49,676,681.33

Cfta_r $ 245,092,067.63

Cfto_r $ 2,119,271.54

Ctsf_r $ -

Cpro_r $ 48,325,522.30

Cfin_r $ 48,325,522.30

CRDTE $ 579,906,269.62

CRDTE/plane $ 193,302,089.21

Figure 14-5: Aircraft Program Life Cycle Costs (Ref. 10)


A summary of the total CRDTE costs are given

in

. Provided in the table are two values for

research and development and costs. The first is the

total expected expenses for the entire program. The

second value is the expected cost per airframe. It was determined that 3 test aircraft would be manufactured (for

flight tests as well as static testing). The CRDTE/plane term reflects the cost of each individual aircraft. A graphical

representation of these results is shown Figure 14-6. Each section represents the percentage of the total CRDTE

expenses.

14.4.2 Acquisition Cost (CACQ)

The importance of a reasonable acquisition cost

cannot be expressed enough as it is extremely important

for the continued success of an aircraft. For the

calculations that were completed, as was previously

stated, it is assumed that 400 units of this aircraft will be

sold. Acquisition cost, in addition to factoring in

program production levels, included the following costs:

Airframe Engineering and Design Cost, Caedm

Airplane Program Production Cost, Capcm

Production Flight Test Operations Cost, Cftom

Cost to Finance Manufacturing, Cfinm

Cost of Manufacturing, CMAN

Cost of Production, CPRO

Caed_r $ 186,367,204.51

Cdst_r $ 49,676,681.33

Cfta_r $ 245,092,067.63

Cfto_r $ 2,119,271.54

Ctsf_r $ -

Cpro_r $ 48,325,522.30

Cfin_r $ 48,325,522.30

CRDTE $ 579,906,269.62

CRDTE/plane $ 193,302,089.21

Caed_m $ 46,305,489.54

Capc_m $ 5,927,297,613.18

Cfto_m $ 9,600,000.00

Cfin_m $ 1,196,640,620.54

CMAN $ 7,179,843,723.27

CPRO $ 1,495,800,775.68

CACQ $ 8,675,644,498.95

CACQ/plane $ 21,689,111.25

Table 14-2: Acquisition Cost Summary

32.1%

8.6%

42.3%

0.4%

0.0% 8.3% 8.3%

Caed_r Cdst_r Cfta_r Cfto_r

Ctsf_r Cpro_r Cfin_r

Figure 14-6: CRDTE Cost Summary

0.5%

68.3%

0.1%

13.8%

17.2%

Caed_m Capc_m Cfto_m

Cfin_m CPRO

Figure 14-7: Acquisition Cost Summary

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