Excerpts From JETT Revise

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

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 (~)
MarketShare,(%)
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
PercentageofMarketShare,(%)
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

σ = +1
σ = -1
σ = -1
σ = +1
2020
Standard
Deviation at σ
= ±1
Design
Point
W/S
Sizing Chart
W/P
W/S
PowerLoading,(W/P)TO(lbf/hp)
Take off Wing Loading (W/S)TO (lbf/ft2
)
)
1998 Example
)
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
SpecificRange,SR(nmi/lb)
Wing Loading, W/S (lbf/ft^2)
AR = 25
AR = 23
AR = 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
Error! Not a valid bookmark
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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