t.a.g.finalproposal_rev(1)
TRANSCRIPT
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Tactical Aeronautic Group
The UR1T
In response to the 2014-2015 AIAA Undergraduate Aircraft Design Competition
California Polytechnic State University, Pomona
Aircraft Design 2014 - 2015
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Team Structure
Adam Ortega
Project Manager
AIAA Member #531155
Dwight Nava
CAD Expert
AIAA Member #530438
Andrea Valdez
Cost Analyst
AIAA Member #531244
George Paguio
Payload Integration
AIAA Member #530448
Tony Ye
Aerodynamics & Mission
Design
AIAA Member #529479
Justin Ellerbee
Takeoff, Landing &
Structures
AIAA Member #531230
Miguel Osorio
Subsystems
AIAA Member #531237
Dong Jin Ryoo
Landing Gear Specialist
AIAA Member #530833
Ramon Navarro
Avionics & Controls
AIAA Member #531242
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Table of Contents Team Structure ................................................................................................................................ 2
List of Figures ................................................................................................................................. 6
List of Tables .................................................................................................................................. 7
Executive Summary ........................................................................................................................ 8
Requirements Compliance Matrix ............................................................................................ 10
UR1T 3-View ............................................................................................................................ 11
1.0 Introduction ............................................................................................................................. 12
1.1 Request for Proposal Summary ........................................................................................... 12
1.2 Mission Profile .................................................................................................................... 12
2.0 Requirements .......................................................................................................................... 13
2.1 RFP Requirements............................................................................................................... 13
2.2 Federal Aviation Regulations .............................................................................................. 14
3.0 Concept of Operations ............................................................................................................ 15
3.1 Manufacturing & Maintenance ........................................................................................... 15
4.0 Historical Aircraft ................................................................................................................... 17
5.0 Initial Aircraft Sizing .............................................................................................................. 17
5.1 Constraint Diagram ............................................................................................................. 17
5.2 Trade Studies ....................................................................................................................... 18
5.3 Detailed Weight Breakdown ............................................................................................... 20
6.0 Fuselage .................................................................................................................................. 21
6.1 Fuselage Layout .................................................................................................................. 21
6.2 Payload Integration ............................................................................................................. 22
6.2.1 Dimensions ................................................................................................................... 22
6.2.2 Payload Placement Configuration Schematics ............................................................. 24
6.2.3 Floor Strength ............................................................................................................... 27
6.2.4 Secondary Design Objectives ....................................................................................... 28
7.0 Wing ........................................................................................................................................ 28
7.1 Wing Configuration & Location ......................................................................................... 28
7.2 Airfoil Selection .................................................................................................................. 30
7.3 High-Lift Devices ................................................................................................................ 32
7.4 Fuel Tanks ........................................................................................................................... 33
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8.0 Propulsion System .................................................................................................................. 34
8.1 Engine Selection .................................................................................................................. 34
8.2 Location ............................................................................................................................... 38
9.0 Landing Gear .......................................................................................................................... 38
9.1 Landing Gear Placement ..................................................................................................... 39
9.2 Landing Gear Loads ............................................................................................................ 41
9.3 Structure .............................................................................................................................. 41
10.0 Tail Sizing ............................................................................................................................. 42
10.1 Horizontal Tail .................................................................................................................. 42
10.2 Vertical Tail....................................................................................................................... 43
11.0 Subsystems ............................................................................................................................ 45
11.1 Auxiliary Power Unit ........................................................................................................ 45
11.2 Electrical System ............................................................................................................... 45
11.3 Avionics ............................................................................................................................ 46
11.3.1 Actuators ..................................................................................................................... 46
11.3.2 Autopilot ..................................................................................................................... 47
11.3.3 Black Boxes ................................................................................................................ 48
11.3.4 Collision Avoidance System ...................................................................................... 48
11.3.5 Communication System .............................................................................................. 49
11.3.6 Weather Radars ........................................................................................................... 49
11.4 Flight Deck ........................................................................................................................ 49
11.5 Fuel System ....................................................................................................................... 50
12.0 Structures & Materials .......................................................................................................... 51
12.1 Structures ........................................................................................................................... 51
12.1.1 Wing Structure ............................................................................................................ 52
13.0 Ground Integration ................................................................................................................ 56
13.1 Loading & Unloading........................................................................................................ 56
14.0 Stability and Control ............................................................................................................. 58
14.1 Wing Contribution............................................................................................................. 59
14.2 Aft tail contribution ........................................................................................................... 59
14.3 Fuselage contribution ........................................................................................................ 59
14.4 Aerodynamic Center ......................................................................................................... 60
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14.5 C.G. Travel ........................................................................................................................ 60
15.0 Aerodynamics ....................................................................................................................... 61
15.1 Drag Polar ......................................................................................................................... 61
15.2 Lift Curve Slope ................................................................................................................ 62
16.0 Performance .......................................................................................................................... 64
16.1 Flight Envelope ................................................................................................................. 64
16.2 Takeoff .............................................................................................................................. 65
16.3 Landing.............................................................................................................................. 69
16.4 Tactical Approach ............................................................................................................. 70
16.5 Payload Range ................................................................................................................... 70
17.0 Cost Estimation ..................................................................................................................... 72
17.1 Acquisition Cost ................................................................................................................ 72
17.2 Operation Cost................................................................................................................... 73
17.3 Life Cycle Cost.................................................................................................................. 74
Conclusion .................................................................................................................................... 76
References ..................................................................................................................................... 78
Acknowledgements ....................................................................................................................... 78
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List of Figures Figure 1: Mission Profile for the Military Transport. ................................................................... 13
Figure 2: Design Point Selection Based on Constraint. ................................................................ 18
Figure 3: Efficient Mach With Respect to Altitude. ..................................................................... 19
Figure 4: Preliminary Aircraft Sizings. ......................................................................................... 20
Figure 5: Cockpit Layout (Taken from Airbus A310 MRTT). ..................................................... 22
Figure 6: Fuselage Schematics...................................................................................................... 23
Figure 7: Cross Section View of Pallet Configuration. ................................................................ 24
Figure 8: Side View of Pallet Configuration. ............................................................................... 25
Figure 9: Top View of Pallet Configuration. ................................................................................ 25
Figure 10: Cross Section View of Bridge Configuration.............................................................. 26
Figure 11: Side View of Bridge Configuration............................................................................. 26
Figure 12: Top View of Bridge Configuration. ............................................................................ 26
Figure 13: Floor Strength Scaffolding. ......................................................................................... 27
Figure 14: Top View of Wing Location. ...................................................................................... 29
Figure 15: Front View of Wing Configuration. ............................................................................ 30
Figure 16: Airfoil Lift Curves. ...................................................................................................... 31
Figure 17: High Lift Devices ........................................................................................................ 33
Figure 18: Fuel Tanks ................................................................................................................... 34
Figure 19: Engine Selection Comparison. .................................................................................... 35
Figure 20: Engine Comparison Selection. .................................................................................... 36
Figure 21: Engine Size Comparison Relative to Thrust. .............................................................. 37
Figure 22: Schematic Bottom View of Landing Gear Placement................................................. 40
Figure 23: Schematic Side View of Landing Gear Placement. .................................................... 40
Figure 24: Notch Chart Diagram. ................................................................................................. 43
Figure 25: UR1T Engine Out Free-Body Diagram....................................................................... 44
Figure 26: APU Location. ............................................................................................................. 45
Figure 27: Control Surface Servo Actuators. ................................................................................ 46
Figure 28: Black Box Data Recorder. ........................................................................................... 48
Figure 29: Inside Flight Deck. ...................................................................................................... 50
Figure 30: Maneuvering and Gust Envelopes. .............................................................................. 51
Figure 36: Cargo Door for Loading 463L Master Pallets. ............................................................ 57
Figure 37: Cargo Door for Loading M104 Wolverine Assault Bridge. ........................................ 58
Figure 38: C.G. Travel. ................................................................................................................. 61
Figure 39: UR1T Drag Polar. ........................................................................................................ 62
Figure 40: Lift Curve Slope of UR1T. .......................................................................................... 63
Figure 41: UR1T Operational Flight Envelope. ........................................................................... 64
Figure 42: Payload vs. Range. ...................................................................................................... 71
Figure 43: Total Program Savings When Learning Curves Applied at 85% and 90%. ................ 74
Figure 44: Life Cycle Cost at 85% Learning Curve. .................................................................... 75
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List of Tables Table 1: RFP Requirements. ......................................................................................................... 14
Table 2: FAR Requirements. ........................................................................................................ 14
Table 3: Aircraft Component Weight Breakdown. ....................................................................... 21
Table 4: Fuselage Specifications. ................................................................................................. 23
Table 5: Payload Specifications. ................................................................................................... 24
Table 6: Auxiliary Unit Payload Considerations. ......................................................................... 28
Table 7: Airfoil Characteristics. .................................................................................................... 31
Table 8: Narrowed Down Engine Selection. ................................................................................ 38
Table 9: Landing Gear Loads. ...................................................................................................... 41
Table 10: UR1T Stall Conditions. ................................................................................................ 51
Table 11: UR1T Operational Limits. ............................................................................................ 64
Table 14: Cruise Segment Specifications. .................................................................................... 71
Table 15: Total Cost of 3 Prototypes at 85% Learning Curve. ..................................................... 73
Table 16:Total Production Cost at 85%, 90%, and 95% Learning Curves................................... 75
Table 17: Total Flyaway Cost of Leading Competitors................................................................ 76
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Executive Summary
UR1T was designed to meet the minimum requirements of the new generation military
transport aircraft expected to be launched by 2030. UR1T is the next generation military
transport aircraft with a span of 211 ft. and a length of 250 ft. The design started with constraint
diagrams that chose the initial values for wing loading and thrust to weight ratio. The thrust to
weight ratio was chosen to be 0.225 and a wing loading of 118 psf. These parameters were then
finalized through an iteration process using a mission program written specifically to evaluate
trade study results. The actual values are 0.28 for the thrust to weight ratio and 122 psf for the
wing loading due to the increase in overall weight from the engine selection, which was required
to provide the necessary thrust. Using empirical formulas for calculating empty weight, the
MTOW for the UR1T was 700,000 lb. With the main parameter acquired, trade studies were
performed to select the optimal airfoil and engine designs for the aircraft. The airfoil that was
chosen was SC(2)-07144 which is a super critical airfoil with a CLmax of 1.77. The wing was
designed to have a span of 211 ft. with 2 flats covering 60 percent of the area of the wing and an
anhedral angle of 5 degrees. The engine selected was the GEnx with a continuous thrust of
72,300 lb with two mounted on each wing. The empennage was designed using the notch chart
technique and simple calculations with considerations of the requirements of the RFP. The
horizontal tail was sized to have a surface area of 1450 ft2
and a span of 65 ft. The vertical tail
was sized to have a surface area of 1460 ft2 and a span of 38 ft. Landing gear was placed based
on load distribution of the payload and other parameters such as tip back and turn around angle.
The fuselage was designed to fit the payload accordingly for both the M104 Assault Bridge and
the Master Pallets. The payload area length across the fuselage is 190 ft. and 23 ft. wide. The
cost was analyzed using learning unit curves at 95%, 90% and 85%. With the 90% learning
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curve the average unit cost came out to be 136 million dollars with a total production cost of 16.3
billion dollars.
UR1T aircraft was a response to AIAA’s RFP for a next generation military transport
aircraft to be launched by 2030. UR1T is by the next generation military transport because of the
fact that it is simple and effective compared to its predecessor, Lockheed’s C-5 Galaxy. UR1T is
lighter than the current military transport, and since our lift to drag ratio is higher the aircraft
burns less fuel. Compared to the current military transport, UR1T loading time are shorter, low
maintenance, and low cost. Since UR1T is a very conventional design, proven technologies are
implemented to reduce the risk of immature technologies causing delays. All of the requirements
have been met and nothing more. Any additional implementation would be unnecessary and
would only result in extra weight, in other words, extra cost in the budget. The UR1T aircraft
will be used throughout the world and it will outperform its previous predecessor not only in
performance but utility as well. There is no doubt that the UR1T will be the best aircraft of its
time and the UR1T is the best design for this program.
UR1T C-5B
MTOW (lb) 700,000 840,000
Range (nm) 1,800 (w/ 300,000lb
payload)
6,300
2,400
Payload (lb) 120,000 263,200
Fuel Capacity (gal) 28,232 51,150
Cost ($) 240,000,000 168,000,000 (in 1987)
Requirements Compliance Matrix
RFP Requirements FAR Requirements
Description Compliance Pg.
#
FAR Description Compliance Pg.
#
6,300 nm unrefueled range with a wartime planned load of
120,000 lb.
YES 71 25-125 Landing Limits YES 67
Maximum payload weight ≥ 300,000 lb. YES 71
Cruise Mach number ≥ 0.60 YES 71
Time to top of climb/climb to initial cruise altitude ≤ 20
min with 205,000 lb.
YES 71
Takeoff field length with maximum payload, and landing
field length with maximum landing weight ≤ 9,000 ft.
YES 68
Takeoff, landing and climb requirements must be met at sea
level in an ISA + 30 C day. Takeoff, and landing
performance should also be shown at ISA+10 C at 10,000’
above MSL.
YES 69
The aircraft shall be able to perform a takeoff, climb to
pattern altitude, conduct pattern flight, and return to base
with one or more engines out immediately after decision
speed. Aircraft with an even number N of engines shall
meet this requirement with any N/2 engine inoperative; if N
is odd then assume N/2 +1 engines inoperative. Indicate the
maximum allowable increase in temperature and altitude
over ISA sea level for which engine(s) out takeoff, as
described here, can be met
YES 44
The aircraft shall be able to perform a tactical approach for
arrivals to bases embedded in combat environments (see
primary design objectives)
YES 70
Internal cargo volume, and corresponding cargo weight
capacity ≥ 44 463L master pallets, or one M104 Wolverine
Heavy Assault Bridge.
YES 24
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UR1T 3-View
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1.0 Introduction
1.1 Request for Proposal Summary
Each year the American Institute of Aeronautics and Astronautics, AIAA, provides a
request for proposal document that requires teams of undergraduate students to design an entire
aircraft system. This year, the AIAA made the objective of the request for proposal to design the
“Next Generation Strategic Airlift Military Transport”, providing major improvements over the
current military transport aircraft fleet, with an anticipated 2030 entry into service (EIS). There
are many aircraft design processes and principles that must be taken into consideration, as well
as a strong understanding of the requirements portrayed in the RFP, in order to meet the given
objective.
1.2 Mission Profile
The RFP specifies a minimum distance that the aircraft must be able to fly. The aircraft
must be able to fly 6,300 nm unrefueled with a planned wartime payload of 120,000 lb., while
minimizing fuel consumption. There are also FARs which are taken into consideration for the
designed mission, such as takeoff and landing requirements, as well as a reserve distance in case
of emergency. A typical mission profile that would be performed by the military aircraft can be
seen in Figure 1. This mission will also incorporate a tactical approach in a wartime
environment.
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Figure 1: Mission Profile for the Military Transport.
2.0 Requirements
The requirements that must be met for the design of an aircraft can come from either the
RFP or Federal Aviation Regulations (FARs). These requirements can give baseline values or
criteria that must be met in order to be certified and considered.
2.1 RFP Requirements
Table 1 shows the main requirements as requested per the RFP. It outlines performance
requirements as well as sizing requirements for the aircraft design.
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Table 1: RFP Requirements.
General Requirements
Requirement Value
Range 6,300 nm @ 120,000# load
Max Payload ≤ 300,000lb
Cruise Mach ≥ .60
Time to Climb ≤ 20 minutes @ 205,000# payload
Takeoff/Landing Field Length ≤ 9000ft @ max payload
Takeoff/Landing Performance
Conditions
Met at SL for ISA +30 C &
Met @ 10,000ft above MSL for ISA +10 C
Mission with Engine
Inoperative
Even – N/2 inoperative engines
Odd – N/2 + 1 inoperative engines
Tactical Approach Aircraft shall be able to perform a tactical approach for arrivals
to bases embedded in combat environments
Cargo Volume 463L Master Pallets – 44 units
M104 Wolverine Heavy Assault Bridge – 1 unit
Climb Speed Limitations Climb Speed < 250kt below 10,000ft
Unit Production 120 units
Entry Into Service By year 2030
2.2 Federal Aviation Regulations
The Federal Aviation Regulations, FARs, are rules or protocols that are created by the
Federal Aviation Administration (FAA) in order to regulate safety and control of commercial
aircraft and airspace. Table 2 shows the most important FAR that were taken into account in the
design of this military transport. This FAR mostly concerned the landing requirements that
would
Table 2: FAR Requirements.
FAR Requirements
Requirement Description
25-125 Landing Limitations
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3.0 Concept of Operations
3.1 Manufacturing & Maintenance
Currently there are no present GE sites that specifically function for regular GEnx
maintenance. At the moment the two largest GE facilities located in Cincinnati, Ohio and Dallas,
Texas perform routine inspections when specialized equipment is necessary. Since these two
sites function as GE operational facilities, there is a need for quicker inspection turnaround time,
especially when out of range of the two MRO sites. Since there is one located in eastern and
western regions, these sites are sufficient for temporary reparation sites. However as the
production is expected to begin as soon as 2016, with an entry into service in 2030, an expansion
of overhaul facilities with the capability to perform sophisticated maintenance is already
underway. The lighter maintenance services are currently tackled by several onsite repair teams
with comparable capabilities to the MRO sites. These repair teams perform line maintenance
services with quick turnaround time, and are sufficient for first quarter of the program, as
manufacturing facilities are built to accommodate the entry of units in to production. Until then
third party partnerships with overhaul sites nationwide provide opportunities to enrich private
industry. TAG Inc partnered with Aerodyne Corp in Stuart Florida, and StandardAero in
Springfield Illinois to officiate them as secondary MRO sites to provide quality service and
support.
TAG Inc maintains a worldwide service network to provide the highest level of
performance and reliability with OEM materials and repairs to fix skin repairs, structural
evaluations, or any supplemental repair at a quick turnaround time, while reducing the overall
cost of operation.
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Operating a worldwide network of dispatch and quick maintenance facilities, any designated
MRO site can provide more than 3,000 rapid repair solutions worldwide. Any supplemental GE
service product or comprehensive evaluation for an ongoing flight line engine maintenance need
can be handled by the onsite repair team, or regional designated support site.
Currently there are no present GE sites that specifically function for regular GEnx
maintenance. At the moment the two largest GE facilities located in Cincinnati, Ohio and Dallas,
Texas perform routine inspections when specialized equipment is necessary. Since these two
sites function as GE operational facilities, there is a need for quicker inspection turnaround time,
especially when out of range of the two MRO sites. Since there is one located in eastern and
western regions, these sites are sufficient for temporary reparation sites. However as the
production is expected to begin as soon as 2016, with an entry into service in 2030, an expansion
of overhaul facilities with the capability to perform sophisticated maintenance is already
underway. The lighter maintenance services are currently tackled by several onsite repair teams
with comparable capabilities to the MRO sites. These repair teams perform line maintenance
services with quick turnaround time, and are sufficient for first quarter of the program, as
manufacturing facilities are built to accommodate the entry of units in to production. Until then
third party partnerships with overhaul sites nationwide provide opportunities to enrich private
industry. TAG Inc. partnered with Aerodyne Corp in Stuart Florida, and StandardAero in
Springfield Illinois to officiate them as secondary MRO sites to provide quality service and
support.
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TAG Inc. maintains a worldwide service network to provide the highest level of performance
and reliability with OEM materials and repairs to fix skin repairs, structural evaluations, or any
supplemental repair at a quick turnaround time, while reducing the overall cost of operation.
Operating a worldwide network of dispatch and quick maintenance facilities, any designated
MRO site can provide more than 3,000 rapid repair solutions worldwide. Any supplemental GE
service product or comprehensive evaluation for an ongoing flight line engine maintenance need
can be handled by the onsite repair team, or regional designated support site.
4.0 Historical Aircraft
The RFP demands the next generation strategic airlift military transport, otherwise, an
aircraft that could replace a Lockheed C-5 Galaxy. This aircraft has a maximum takeoff weight
of 840,000 lb. with a maximum range of 2,000 nm. The specifications of this aircraft was a good
starting point for the initial parameters of the aircraft design. A further discussion on the
comparison of the UR1T and previous transports will include the overall improvements of the
new design.
5.0 Initial Aircraft Sizing
In order to preliminarily size the aircraft, certain parameters (such as T/W, W/S, wing
AR, etc.) were estimated based on meeting the requirements of the RFP mission. These
parameters set the initial design point which is discussed in the next section.
5.1 Constraint Diagram
The constraints shown in Figure 1 were created using kinematic equations. The initial
design point was chosen based on having the lowest thrust to weight ratio, and the lowest wing
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loading. Having the lowest thrust to weight ratio allows the aircraft to have smaller engines,
reducing fuel consumption and overall takeoff weight. However, when actually doing the
calculations, the thrust to weight ratio was not 0.225 due to the engines selected, resulting in a
0.28 thrust to weight ratio. Due to the higher weight from the engines, more fuel was required,
such that the wing loading was also increase from the initial design point of 118 psf to 122 psf.
Figure 2: Design Point Selection Based on Constraint.
5.2 Trade Studies
A mission program was developed in order to vary certain parameters in the aircraft’s
flight in order to find the optimum conditions for the aircraft. The first conditions that were
varied were Mach number and altitude for flight. Figure 3 shows a visual representation of the
best conditions for the aircraft. Based on the initial design point, wing loading conditions
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constrained the area of which would be best fit for the plane. It was decided that Mach 0.75 at an
altitude of 30,000 ft. was the most optimum with respect to maximum takeoff weight.
Figure 3: Efficient Mach With Respect to Altitude.
The next plot describes the conditions for the initial sizing of the aircraft. Again, the wing
loading of the initial design point was a constraining factor, in which the parameters chosen were
close to the wing loading. The most efficient conditions for the aircraft was found using an
aspect ratio of 7.75 with a wing area of 5,750 ft2. Further iterations were performed in order to
verify the chosen parameters for the aircraft.
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Figure 4: Preliminary Aircraft Sizings.
5.3 Detailed Weight Breakdown
A very important aspect of any aircraft design is the weight breakdown. Table 3 shows
the empty weight breakdown of the components of the aircraft. These values were found using
empirical formulas found in Carichner’s Aircraft Design book which were verified using similar
aircraft specifications. These empty weights along with payload and fuel will be used in order to
find the center of gravity of the aircraft at various loading conditions.
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Table 3: Aircraft Component Weight Breakdown.
Element Weight (lb..) Element Weight (lb..)
Wing 88,346 Starting System 4.266
Fuselage 74,866 Surface Controls 13,086
Propulsion 44,390 Instruments 1,208
Landing Gear 35,359 Electrical 9,846
Fuel System 9,623 Furnish 8,443
Horizontal Tail 16,814 A/C + Icing 5,185
Vertical Tail 13,691 Avionics 1,700
Engine Controls 193 Empty Weight 327,034
6.0 Fuselage
6.1 Fuselage Layout
The cockpit was designed to fit two people: the Captain and First Officer for each
mission. To accommodate this mission crew, the cockpit dimensions were set to 25 feet in length
and a maximum width of 25 feet at the entrance wall. In designing the cockpit, FAR guidelines
were taken into account in adding a galley, lavatory, and ample spacing between instrument
panels. In lieu of designing a cockpit from scratch, the TAG UR1T's cockpit was based and
designed to follow the configuration of the Airbus A310 MRTT's cockpit as seen in Figure 5
below:
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Figure 5: Cockpit Layout (Taken from Airbus A310 MRTT).
*Figure taken from http://www.airforce-technology.com/projects/mrtt/mrtt7.html
6.2 Payload Integration
Being a transport aircraft, payload capability is one of the integral portions of the design
process for the TAG UR1T. Specifically, the fuselage and cargo bay area was designed to meet
the requirement of "internal cargo volume, and corresponding cargo weight capacity, shall be no
less than 44 463L Master Pallets, or one M104 Wolverine Heavy Assault Bridge". These were
designed with the objective of minimizing turn-around time for efficient missions.
6.2.1 Dimensions
The TAG UR1T fuselage was designed around a conventional cylindrical shape.
Specifications of the fuselage are shown in Table 5 below:
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Figure 6: Fuselage Schematics.
Table 4: Fuselage Specifications.
Length (ft.) 190
Max Width (ft.) 25
Max Usable Height (ft.) 15
Fuselage sizing was determined due to a couple key design decisions. Traditionally,
cargo aircraft are sized to have longer and thinner fuselages to minimize drag, increase takeoff
and landing performance, and minimize tail sizes. The TAG UR1T width was designed to hold
two rows of the 463L Master Pallets lengthwise in order to achieve the previously stated goal.
Another design decision was to only load from the back of the aircraft through the cargo door.
This was done to reduce loading time of the M104 Wolverine Assault Bridge due to the CG of
the aircraft being closer to the tail end. It also reduces the quantity of people required to load the
aircraft along with one less forklift required when loading the 463L Master Pallets.
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Being a military cargo aircraft, having the fuselage designed around proper integration of
payload outlined in the RFP was paramount. The two different types of payload are detailed in
Table 5 below:
Table 5: Payload Specifications.
Specifications 463L Master Pallets M104 Wolverine Assault Bridge
Length (ft.) 9 44
Width (ft.) 7.34 11.42
Height (ft.) Negligible* 13
# Units Required 44 1
*Pallet height is very thin, but can be assumed no larger than 13ft due to Bridge height. Average
Pallet cargo height is 8ft.
6.2.2 Payload Placement Configuration Schematics
To do accomplish this, the 463L Master Pallets are designed to be placed in the following
configurations:
Figure 7: Cross Section View of Pallet Configuration.
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Figure 8: Side View of Pallet Configuration.
Figure 9: Top View of Pallet Configuration.
Additionally, the M104 Wolverine Assault Bridge is designed to be placed as shown
below:
26
Figure 10: Cross Section View of Bridge Configuration.
Figure 11: Side View of Bridge Configuration.
Figure 12: Top View of Bridge Configuration.
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6.2.3 Floor Strength
To accommodate for the large amount of cargo required to be transported by the UR1T
aircraft, proper floor strengthening needed to be installed in accordance to the types of payload
needing to be integrated. To determine the constraining factor of the flooring strength of the
cargo area, stress analysis was done with treating the M104 Wolverine Assault Bridge as a point
load, and the 463L Master Pallets as a distributed load across the floor area. The constraining
load was found to be the M104 Wolverine Assault Bridge point load due to the large weight of
the bridge in comparison to its volume. To be thorough, the floor was strengthened as shown in
Figure 13 below with I-beam scaffolding:
Figure 13: Floor Strength Scaffolding.
There is scaffolding running lengthwise down the fuselage underneath where the 463L
Master Pallets would be place in the configurations shown above, along with extra beams
underneath where the M104 Wolverine Assault bridge would be placed to deal with the extra
point load.
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6.2.4 Secondary Design Objectives
Auxiliary payload units were considered via the AIAA RFP to be able to fit into the
UR1T to allow for more versatile transport uses. Units that were considered and the number of
units able to fit into the UR1T are shown in Table 6 below:
Table 6: Auxiliary Unit Payload Considerations.
M1A Abrams Tank 5
M2/M3 Bradley Infantry Vehicle 8
Apache Helicopter (w/ rotor blades unattached) 6
It should be noted that these values are purely the amount of units of each auxiliary
payload type that can fit into the UR1T volumetrically: the weight of the payload is not being
considered into this calculation. The floor strength and aerodynamic data are not designed to
hold this amount of added weight, but are shown here as an option to the customer if desired
later in the project lifetime.
7.0 Wing
7.1 Wing Configuration & Location
The air vehicle proposed will have a high mounted wing. An anhedral of 5° was added to
the wing. The high mounted wing is preferable for military applications due to runways with
debris. With the wing mounted higher, the risk of foreign object damage to the engines is
reduced. Since one of the design objective was for efficient loading and unloading, having a
cargo floor close to the ground would be advantageous for easy loading. With the wing mounted
above the fuselage, the landing gear is mounted in the fuselage making it shorter and more
29
compact and yet still maintaining ground clearance for the engines and wing tips. The wing
structure is discussed in detail in the structures section. The aerodynamic center is located 120ft
from the nose of the aircraft. This location was chosen because it was behind the center of
gravity and is far forward enough to minimize the size of the horizontal tail to rotate the aircraft
during takeoff.
Figure 14: Top View of Wing Location.
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Figure 15: Front View of Wing Configuration.
7.2 Airfoil Selection
Three supercritical airfoils and one conventional airfoil were considered. The
supercritical airfoils are the NASA SC(2)-0714, SC(2)-0614 and SC(2)-0414. The conventional
airfoil was the NACA 23012, and this airfoil was considered because of its high lift coefficient
and low drag at cruise. The airfoil lift curves are shown in Figure 16 below, generated from
airfoiltools.com.
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Figure 16: Airfoil Lift Curves.
All four airfoils have a stall angle of around 15° and the airfoil with the highest Clmax of
1.77 is the SC(2)-0714 supercritical airfoil. Another parameter considered in airfoil selection was
the critical Mach number. According to NACA L-357 the critical Mach number for 23012 airfoil
is 0.645. Supercritical airfoils are designed to have delayed critical Mach numbers at around 0.7.
Since the cruise speed is within the transonic regime, a higher critical Mach number would be
desirable, meaning less sweep of the wing which also means lower structural weight and better
low speed lift characteristics. Table 7 compares the four airfoils side by side.
Table 7: Airfoil Characteristics.
SC(2)-0714 SC(2)-0614 SC(2)-0414 NACA 23012
Clmax 1.77 1.75 1.65 1.51
Stall Angle (deg.) 15.5 15.5 15.6 15.3
Critical Mach 0.75 0.75 0.75 0.645
32
The airfoil selected for the design was the SC(2)-0714 due to it having the highest lift and
a high critical Mach number.
7.3 High-Lift Devices
To meet takeoff, and landing requirements the lift produced from the plain wing is
insufficient. A high lift device system was integrated to increase lift during takeoff and landing.
Analysis was done for the two required atmospheric conditions for takeoff and landing and the
ISA +10ºC at 10,000ft above mean sea level was determined to be the constraining condition.
The lower density makes landing within 9,000ft difficult and required a lift coefficient of 2.7.
Since landing at this altitude was the most demanding on lift, 2.7 is the maximum lift coefficient
of the whole vehicle.
To achieve a lift coefficient of 2.7, a leading edge slat and a double slotted flap on the
trailing edge was employed. The leading edge slat will run along almost the entire span. The flap
will take up 30% of the chord and the projected area of the flap will be 60% of the wetted area of
the wing. The meaning of projective area is shown in Figure 17 along with the high lift devices
on the wing. The plain wing gives a maximum lift coefficient of 1.55 and the high lift devices
add 1.15 giving a total of 2.7.
33
Figure 17: High Lift Devices
7.4 Fuel Tanks
Since the airfoil selected has a thickness ratio of 0.14, fuel tanks that are large enough can
easily fit inside the wing structure alone. Not having fuel tanks inside of the fuselage allows for
more room and clearance for larger cargo, such as the M104 Assault Bridge. The fuel that is
being stored is JP-8 and 28,232 gallons of fuel is required to perform the mission. The two fuel
tanks are located inside the starboard and port wing structure. Integral tanks, also known as “wet
wing”, are used in order to reduce weight. With a wet wing, the fuel tank portion of the wing
structure is simply sealed off rather than having a hard fuel tank placed inside. Figure 18
illustrates the fuel tanks which are highlighted in orange. The fuel tanks are 22 feet wide at the
wing root, 15 feet wide at the further edge, 45 feet long and 2.4 feet deep.
34
Figure 18: Fuel Tanks
8.0 Propulsion System
8.1 Engine Selection
To determine which engine would suffice for the Military Transport, several different
engines were compared side by side to determine the one best fit for the project design envelope.
The driving requirements was that the minimum cruise speed must exceed Mach .6, as well as
the take-off and landing field length requirement of no more than 9000ft. With this in mind a
spreadsheet was generated that could compare the standard sea level thrusts, specific fuel
consumptions and weights possible for each engine. Figure 19 displays the 24 different engines
considered plotted by each respective dry weight against its sea level thrust. Notice the three best
contenders included the GEnx, Trent 1000, and PW 4000. These three engines allowed the
design to deliver at least a cruise Mach number of 0.6, for the conventional four engine design.
The chosen engine sufficient for the four engine design, would have to consider the
weight of the four engines plus the maximum payload weight could not be any lower than
300,000lb, and 120,000lb during wartime planned load. This meant that the engine should
deliver not necessarily the highest thrust possible but enough to deliver the mission expectation,
35
with minimum weight. This allowed us to narrow down the search to engines with thrust levels
between 55,000-65,000 lb. thrust.
Figure 19: Engine Selection Comparison.
According to our estimations engine weight must not exceed 14,000 lb. per engine for the
two engines mounted on each wing. This weight agreed with the maximum loads the wing could
be support in the rate of climb requirement of initial cruise altitude in 20 minutes, with 205,000lb
payload. It quickly became important not to disrupt the cg location, by keeping the wing weight
and dimensions fixed. To do so a longer engine would keep the wind load distribution as
symmetrical along the wing as possible. The diameter of the engine should be largest possible to
optimize the turbine efficiency. Though these parameters are standard with each engine, these
parameters were important to solve for the engine location.
PW 4000
Trent 1000
GEnx
y = 0.4054x0.9255
0
2000
4000
6000
8000
10000
12000
14000
16000
18000
20000
0 20000 40000 60000 80000 100000
Dry
We
igh
t
SLS Thrust
Engine Dry Weight
36
Figure 20 shows the different engine’s output thrust versus the diameter of each engine.
Notice that the GEnx engine has the largest diameter to thrust ratio and length to thrust ratio,
making it the best contender with the least cg shift on the four engine design. The closer to the
trend line the less cg shift affect the engine would add to the design. This is what separated the
PW4000 engine series and the Trent 1000 from the GEnx.
Figure 20: Engine Comparison Selection.
PW 4000
Trent 1000
GEnx
y = 1.2535x0.3954
0
20
40
60
80
100
120
140
160
0 20000 40000 60000 80000 100000
Dia
me
ter
(in
)
SLSThrust
Max. Diameter vs. Thrust
37
Figure 21: Engine Size Comparison Relative to Thrust.
The final deciding factor for the engine selection after the top three contenders were
selected were to consider the specific fuel consumption at standard sea level for each engine,
since each engine would have to meet the 6300 nautical mile unrefueled range wartime
requirement with a 300,000lb payload. Table 8 exhibits the most important characteristics of the
engines such as specific fuel consumption, thrust, weight, and cost, where red is the worst option,
and green is the engine with the best value for that category. Notice the GEnx engine contains
the greenest options, and the fewest red options.
PW 4000
Trent 1000
GEnx
y = 1.6826x0.4284
0
50
100
150
200
250
300
350
0 20000 40000 60000 80000 100000
Engine Length vs. Thrust
38
Table 8: Narrowed Down Engine Selection.
Key Factors PW 4000 Trent 1000 GEnx
Continuous Thrust 52,000 lb. 58,000 lb. 72,300 lb.
Weight 9,570 lb. 12,700 lb. 12,800 lb.
T/W 5.4 4.6 5.65
SFCSL .361 .35 .348
Unit Cost 15 M 16 M 13 M
8.2 Location
The engine locations are placed at 50ft and 80ft relative to the centerline of the aircraft.
This distance of 50ft was selected in order to avoid flow effects from the fuselage disturbing the
inlet air of the engines. The 80ft distance was to provide adequate space between the engines
while still being far from the wing tips. The closer to the centerline the less of an effect the
engine will have on yawing the aircraft
9.0 Landing Gear
Landing gear design is often lightly considered, yet important, part of the overall aircraft
design, taking into consideration its configuration, performance, and weight. The basic
requirement of a landing gear is to withstand the maximum load provided by the aircraft,
including the stress from takeoff, taxi, and landing. There are also the gear’s placement, weight,
and structural components to consider in its design process.
39
9.1 Landing Gear Placement
The first and most important requirement of the landing gear to consider is its placement. It is
positioned relative to the plane’s center of gravity, and is a variable to consider in finding a
proper weight distribution between the nose and main gear. According to Nicolai & Carichner’s
book, a safe and common assumption of the weight distributed between the nose and main
landing gear is twenty percent to eighty percent respectively –. With the assumption of the loads
being static, the proper weight percentage on the nose landing gear can be found out by
𝑊𝑁𝐺 =𝑊𝑇𝑂∗|𝑋𝑀𝐺−𝑋𝑁𝐺|
𝑋𝑛𝑔
Where WTO is the maximum takeoff weight of the aircraft, XMG is the distance of the main gear
from the nose, and XNG is the distance of the nose gear from the main. The calculation for the
main gear is a similar but opposite of the nose gear as seen in the equation
𝑊𝑀𝐺 =𝑊𝑇𝑂∗|𝑋𝑁𝐺−𝑋𝑀𝐺|
𝑋𝑀𝐺
A visual representation of these equations can be seen in Figure 22.
40
Figure 22: Schematic Bottom View of Landing Gear Placement.
Figure 23: Schematic Side View of Landing Gear Placement.
Another factor that will affect the placement of the landing gears is the tip-back angle –
the angle of rotation of the aircraft. The angle calculated considers the necessary angle of attack
41
to achieve flight and the root angle of twist in order for the rear end of the plane to scrape the
ground during takeoff. This can be altered by not only the placement of the main landing gear,
but also its height. Considering the RFP requirement of a faster load time, the minimum height
requirement was considered. Figure 23 shows a visual representation of how the height of the
landing gear can affect the tip-back angle of the aircraft.
9.2 Landing Gear Loads
After carefully considering all the requirements needed for the landing gear placement
Table 9 shows the specification of the landing gear. The equations for these calculations may be
found in Nicolai & Carichner.
Table 9: Landing Gear Loads.
% Weight Weight (lb.) Height (ft.) Width (ft.) From Nose (ft.)
Main Gear
18.6
130,259
10
21.5
25
Nose Gear
81.4
568,239
10
-
150.7
9.3 Structure
The main consideration of the structural design of the landing gear was the strut,
deployment and retraction of the landing gear, and the tires. Due to the lack of information –
current designs employed on aircrafts were kept confidential due to proprietary reasons – the
design considerations were compared to and chosen from currently existing aircrafts. The main
landing gear follows the Boeing 777 design, where each strut supports 4 wheels – 2 duals in
tandem – and has a total of 4 struts and 16 wheels. The nose landing gear consists of the same
design with the exception of it being steerable compared to the locked main gear and only using
42
1 strut with 4 wheels. Both the main and nose gear will be deployed in a parallel direction to the
fuselage by rotating down from the nose to tail. This is done to decrease the added stress on the
landing gears created by drag during landing. The tires used for the landing gear will also be the
same as the Good Year Aviation tires used on the Boeing 777. It fits the requirement of the total
load each tire can handle as well as the landing speed requirement of 178 knots (tolerates up to
190 knots).
10.0 Tail Sizing
10.1 Horizontal Tail
The horizontal tail is one of the main aspects when considering whether or not an aircraft
is stable. In order to size the horizontal tail, preliminary estimations are made based on the
geometry of the aircraft. A volume coefficient can be obtained, resulting in a reference area. For
further analysis, Figure 22 shows a more detailed sizing for the horizontal tail. The landing flare,
rear stability limit, and nosewheel liftoff are the three constraints for the tail sizing. These
constraints were found base on empirical equations as well as kinematic equations. Once plotted,
limits are made, in which the center of gravity travel must reside within those limits shown in the
figure. The center of gravity moves about ten percent with respect to the leading edge MAC. The
horizontal tail volume coefficient of 0.8 was the minimum size for the horizontal tail. Further
calculations based on the geometry of the wing lead to a 1,460 ft2 area for the horizontal tail.
43
Figure 24: Notch Chart Diagram.
10.2 Vertical Tail
The vertical tail is sized based on a two-engine out scenario. Under this condition, a
moment is created about the center of gravity due to the unbalanced thrust forces. The vertical
tail is used to counteract this moment so that the plane can still be controlled while in flight. In
order for there to be control, the vertical tail must be the right size so that a large enough tail lift
force is created when the rudder is deflected. The force body diagram for this calculation is
shown in Figure 25.
44
Figure 25.
Figure 25: UR1T Engine Out Free-Body Diagram.
First the lift force must be solved by summing the moments about the center of gravity.
𝛴𝑀 = 𝑇 ∗ 50 + 𝑇 ∗ 80 − 𝐿 ∗ 100 = 0
Where the thrust of each engine is 62,000 lb.; therefore the lift force of the vertical tail is 80,600
lb.. Using the Xfoil program, the maximum CLvt was calculated to be 0.8 with trailing edge fully
deflected. This can then be used to solve surface area of the vertical tail.
𝐶𝐿 =𝐿
𝑞∗𝑆𝑣𝑡 𝑆𝑣𝑡 = 1,530 𝑓𝑡2
45
11.0 Subsystems
11.1 Auxiliary Power Unit
The APU chosen for the UR1T aircraft was Pratt and Whitney’s Aeropower PW 980. It is
a two shaft gas turbine engine that provides bleed air for cabin conditioning and main engine
starting. It also provides electrical power from two gearbox-mounted, 120 kilovolt amp (kVA)
generators. This APU was selected since it fit onto the UR1T wide body, which has been used
for wide body aircraft such as the Airbus A380. An important factor in selecting one power unit
was the need to deliver in-flight back up power. The APU’s location will be in the rear end of the
aircraft.
Figure 26: APU Location.
11.2 Electrical System
The UR1T electrical system basically consists of batteries, generators, transformer-
rectifiers (TR’s), electrical controls, circuit breakers, and cables. It will contain four engine-
driven AC generators, each capable of providing sufficient electricity for all the aircraft’s
systems. Each main generator provides a maximum of 110 kVA. The TR’s are used to convert
46
alternating current to direct current. A DC power distribution system will be utilized for battery
and emergency operations. In addition, the UR1T will employ a split-parallel power distribution
system.
11.3 Avionics
11.3.1 Actuators
The control surface actuators will include multiple sensors in the control loop such as
attitude gyros, rate gyros, altimeters, and velocity. The transfer functions for most sensors will be
approximated by the “k” gain. The control surface servo actuators will be hydraulic due to the
high loads that the aerodynamic control surfaces will be exposed to and it can be seen in Figure
27.
Figure 27: Control Surface Servo Actuators.
47
11.3.2 Autopilot
The autopilot will be the most advanced system in the aircraft; it will require minimum
input from the pilot in the cockpit. There are going to be multiple flight computers for
controlling independent subsystems which ultimately will increase two things; processing power
and redundancy in the system.
The first and most reliable autopilot system included in the controls system package is the
displacement autopilot. The displacement autopilot integrates a pitch displacement autopilot, roll
attitude autopilot, altitude hold control system, and a velocity hold control system. The pitch
displacement autopilot works with a vertical gyro that senses the pitch angle, compares it to the
input/desired pitch angle and the error is utilized to produce the required elevator movement to
correct itself. The roll attitude autopilot works in a similar manner but it utilizes an attitude gyro
to measure the roll angle and the ailerons to attain the desired angle. Altitude of the aircraft will
be held constant with the use of an altitude hold autopilot; this type of autopilot assumes constant
velocity and neglects any lateral dynamic effects so that the only motion available is the vertical
plane. Last but not least there is also a velocity hold control system which functions with the
propulsion system. The main purpose of this system is done by regulating the throttle of the
engines to obtain the desired flight speed.
In addition to the autopilot, UR1T will be equipped with the most advanced Stability
Augmentation System (SAS) to provide stability when the flight conditions are not optimum.
This system will have dampers in the three main aircraft axes; roll, pitch, and yaw. It will also be
capable of maintaining stability when encountering really high gusts which cause the structure to
flutter.
48
11.3.3 Black Boxes
This aircraft will be equipped with two data recorders, shown in Figure 28, and they will
not only record audio but also video of all the activity happening in the cockpit as well as the
flight data while in operation. In case of a fatal failure, this would allow engineers to determine
the cause of the accident more accurately and in a faster manner.
Figure 28: Black Box Data Recorder.
11.3.4 Collision Avoidance System
With the tremendous increase in air traffic, the probabilities of a collision increase
exponentially. To account for that problem the UR1T is equipped with a collision avoidance
system. The way this system works is using GPS coordinates and a long range radar; when the
computer senses the possibility of a collision the aircraft will immediately start to change its
altitude and heading angle. The long range radar and high accuracy differential GPS will be able
to detect the collision several miles ahead which will allow more than enough time for the
aircraft to avoid it.
49
11.3.5 Communication System
This aircraft will be equipped with the best communication system made by 2019. It will
have triple redundancy in case weather conditions get too rough or the signal fails and there will
be several ground stations around the globe to transmit telemetry and check the status of the
aircraft at all times.
11.3.6 Weather Radars
To anticipate bad weather conditions such as thunder storms and heavy turbulence areas,
UR1T will have weather radars. As soon as an abnormal condition is detected the waypoints will
be modified in order to go around that area to accomplish a safer and more fuel efficient flight.
The avionics system of this aircraft will include all improved and updated systems at the
2019 technology freeze date, which include systems such as aircraft management/warning, flight
controls, weather monitoring, collision detection, black boxes, communications, and navigation.
11.4 Flight Deck
The URT1 aircraft will offer the most technological flight deck from the aircraft industry
with its top of the line high definition displays to give the pilot a better view of the aircraft’s
flight status. Also, by making all the electronics smaller the space available for the pilots will be
increased, giving the pilots a more pleasant and less tiring flight. This flight deck will also
incorporate highly ergonomic seats which will make the pilot not want to get off the plane after a
bringing it to a safe landing. The sketch of the flight deck is shown in Figure 29 and it is in
compliance with the FARs 25.771 and 25.781.
50
Figure 29: Inside Flight Deck.
11.5 Fuel System
The UR1T has a fuel capacity of 235,500 lb. Two integral tanks will be located inside the
wings. A very important consideration includes actively regulating the fuel during operation, so
as to minimize the deflection of control surfaces. By using transfer pumps and valves throughout
the system, it will assist in maintaining an appropriate aircraft center of gravity.
51
12.0 Structures & Materials
12.1 Structures
As a starting point for structural analysis, the loads on the UR1T aircraft are required.
The V-n diagram shown in Figure 30 was constructed to show the anticipated load factors due to
maneuver speeds and gust loads.
Figure 30: Maneuvering and Gust Envelopes.
Table 10 shows the speeds and loads at different conditions of stall, maneuver, cruise and dive.
Table 10: UR1T Stall Conditions.
Condition Value
Stall: Vs 176 keas
Maneuver: VA 310 keas
Max Gust, VB 330 keas
Cruise, VC 441 keas
-1.5
-1
-0.5
0
0.5
1
1.5
2
2.5
3
3.5
0 100 200 300 400 500 600Speed, keas
Load
Fact
or,
n
Vs
VA VC VD
52
Dive, VD 552 keas
Gust load factors, nC +2.6, -0.6
Load limit factors, n +3.0, -1.0
12.1.1 Wing Structure
For a jet transport, the loading limit factors are defined as 3.0 and -1, according to FAR
Part 25. The velocity for maximum gust intensity of 330 knots produced the load factor of 2.6. It
should be noted that this speed, VB, is greater than the maneuver speed, but less than the cruise
speed. Also, the loading factors at cruise would be used to obtain the aircraft’s primary loads.
The reason for this is the load factors are highest at this condition, thus they will be used to
design the aircraft’s main structural components.
.
53
Figure 31: UR1T Wing Shear Distribution
From the shear forces, bending moments can be solved as well:
54
Figure 32: UR1T Wing Bending Moment Distribution
These forces and moments diagrams are used in order to determine the maximum forces
on the wing. Where maximum bending moment is 12,000,000 ft lbs, and maximum shear force
is 221,000lbs.
Based upon AIAA article 2004 – 1624 AC Struct Layout, it was decided to use 2 spars.
One spar is placed at 30% of the chord length, and the other at 70%. This decision was based on
aircraft of similar size and weight as seen in.
55
Figure 33: Spar Location of Boeing Aircraft
Because most of the aerodynamic loads are near quarter chord of the wing, the front spar
will see significantly more of the loads and will have to be sized accordingly. Using the bending
moment equation, the area of the web can be solved for, and therefore the thickness. The
thickness of the wing determined the height of the web, therefore thickness can be solved for
after area is calculated.
After finding the bending moment and selecting a material:
𝜎𝑦 =𝑀
𝐴𝑧
ℎ𝑓𝑟𝑜𝑛𝑡 = 7 ft
ℎ𝑟𝑒𝑎𝑟 = 4 ft
𝐴𝑓𝑟𝑜𝑛𝑡 = 0.80 ft2 𝐴𝑟𝑒𝑎𝑟 = 0.313 ft2
𝑡𝑓𝑟𝑜𝑛𝑡 = 0.229 ft
𝑡𝑟𝑒𝑎𝑟 = 0.157 ft
56
Figure 29 shows these sizings:
Figure 34: Spar Web Sizing
The placement of wing ribs, and fuselage frames were also selected based on AIAA article 2004
– 1624. Ribs are placed every 2 feet in the span direction, and frames are placed every 20
inches. An FEA model would provide a better representation of the internal loads that the
structure will be placed under.
13.0 Ground Integration
13.1 Loading & Unloading
Reducing loading and unloading time of the required payload is one of the main design
objectives going into the UR1T’s design. To do this, a rolling conveyor belt floor is installed into
the cargo area floor in the same placement as shown to match the configuration drawings for the
57
463L Master Pallets. Loading protocol differs depending on the payload being taken on a given
mission. Unloading of the same configuration payload would be similar to the loading times.
Loading of the 463L Master Pallets requires a minimum one forklift able to carry the
463L Master Pallet and corresponding load on top of it onto the cargo door set horizontally 8.5ft
above the ground as shown in Figure 36 below:
Figure 35: Cargo Door for Loading 463L Master Pallets
Upon lifting the 463L Master Pallets onto the cargo door, they are then pushed into their
corresponding places by the cargo loading crew manually. Assuming the average human walking
speed of 4.55 ft./s and an average pallet to cargo door fork lift speed of 2 pallets placed every 30
seconds, it takes a total of 19 minutes to fully load the 44 units of the 463L Master Pallets into
the configuration shown in the Placement Configuration section.
Loading the M104 Wolverine Assault Bridge requires the cargo door to be lowered as a
ramp onto the floor as shown in Figure 37 below:
58
Figure 36: Cargo Door for Loading M104 Wolverine Assault Bridge.
The M104 Wolverine Assault Bridge is then required to drive up the ramp and fixed into
the configuration required shown in the Placement Configuration section. Assuming a drive
loading speed of 5.87 ft. /s, the load time of the M104 Wolverine Assault Bridge is estimated at
20 seconds.
14.0 Stability and Control
The stability of the aircraft is a crucial aspect of the design process. There are various
methods of evaluating an aircraft’s stability, the one utilized in this case is outlined in Nelson’s
“Flight Stability and Automatic Control”, and it is denoted on the following paragraphs.
In order to obtain static longitudinal stability the aircraft’s pitching moment slope must be
negative and shall comply with the following criterion:
𝐶𝑚𝛼 = 𝑑𝐶𝑚
𝑑𝛼=
𝑑𝐶𝑚
𝑑𝐶𝐿 𝑑𝐶𝐿
𝑑𝛼
However, the pitching moment of the aircraft is determined by all the other aerodynamic
components such as wing, tail, fuselage, and engine therefore it is necessary to compute each
individually.
59
14.1 Wing Contribution
The contribution of the wing to the pitching moment of the aircraft can be summarized to
these two equations, which have been adapted to the static stability conditions:
𝐶𝑚0𝑤= 𝐶𝑚𝑎𝑐𝑤 + 𝐶𝐿0𝑤
(𝑥𝑐𝑔
𝑐−
𝑥𝑎𝑐
𝑐)
𝐶𝑚𝛼𝑤 = 𝐶𝐿𝛼𝑤
(𝑥𝑐𝑔
𝑐−
𝑥𝑎𝑐
𝑐)
14.2 Aft tail contribution
Once the total lift generated by the wing is calculated, a summation of the moments is done
with the assumption of small angles of attack. Then a new term is introduced to the equation;
horizontal tail volume ratio, and the expressions that yield the pitching moment contribution
from the tail are:
𝐶𝑚0𝑡= 𝜂𝑉𝐻𝐶𝐿𝛼𝑡
(휀0 + 𝚤𝑤 − 𝚤𝑡)
𝐶𝑚𝛼𝑡= −𝜂𝑉𝐻𝐶𝐿𝛼𝑡
(1 −𝑑휀
𝑑𝛼)
14.3 Fuselage contribution
One would imagine that the fuselage just has a single role in the aircraft design; carry the cargo
and/or passenger. This is true but it also has some effect on an airplane’s aerodynamic properties.
It affects the pitching moment in a significant way such that we need to account for it when
doing the calculations. The following equation gives us the fuselage contribution to the pitching
moment:
𝐶𝑚𝛼𝑓=
1
36.5 ∗ 𝑆 ∗ 𝑐∑ 𝑊𝑓
2
𝑥=𝑙𝑓
𝑥=0
𝛿휀𝑢
𝛿𝛼Δ𝑥
60
.
14.4 Aerodynamic Center
One of the most important calculations is the aerodynamic center because this is where the
pitching moment coefficient is independent of angle of attack. Since not all the data was readily
available we had to make some assumptions.
��𝑎𝑐 =
��𝑎𝑐,𝑤𝑏 + 𝐶𝐿,𝛼𝐻
𝐶𝐿,𝛼𝑊𝐵𝑉𝜂𝐻
𝑆𝐻𝑇
𝑆𝑅𝑒𝑓��𝑎𝑐,𝐻 (1 −
𝑑휀𝑑𝛼
)
1 +𝐶𝐿,𝛼𝐻
𝐶𝐿,𝛼𝑊𝐵𝑉𝜂𝐻
𝑆𝐻𝑇
𝑆𝑅𝑒𝑓(1 −
𝑑휀𝑑𝛼
)
14.5 C.G. Travel
In order to estimate the center of gravity travel, a weight breakdown was developed. This
was done for 5 different scenarios. The five scenarios simulated were empty, empty operational
weight, zero fuel, zero payloads, and maximum takeoff gross weight (fully loaded).
As mentioned previously, in order to achieve static longitudinal stability, the aerodynamic
center must be located aft of the CG at all times. This was accomplished by shifting components
and even the entire wing aft or forward from the cg. This process was first done with no weight
and once equilibrium was reached then the fuel and other weights were added to simulate the
remaining case scenarios. The following Figure 38 is a plot of the center of gravity travel with
respect to the distance to the MAC.
61
Figure 37: C.G. Travel.
15.0 Aerodynamics
15.1 Drag Polar
The drag polar shown in Figure 39 was generated using the parabolic algebraic equation
of zero-lift drag coefficient and induced drag. The zero lift drag coefficient used was 0.015. This
value is a standard value when it comes to predicting drag at the conceptual level. The lift to drag
ratio at the cruise point is 17.8. This point is very close to the maximum lift to drag point that lies
tangent to the drag polar.
62
Figure 38: UR1T Drag Polar.
15.2 Lift Curve Slope
The lift curve slope for the UR1T air vehicle is shown in Figure 40 below. This plot was
generated using two dimensional airfoil data, from X-FOIL along with the methods shown in
Chapter 9 of Nicolai & Carichner. There are two curves on this plot, the plain wing and Landing
and Takeoff. The plain wing curve is just the lift generated from the wing itself with no high lift
devices deployed. The Landing and Takeoff curve is with all high lift devices deployed. With all
the high-lift devices deployed at their maximum deflection, takeoff distance and landing speed
0
0.2
0.4
0.6
0.8
1
1.2
0 0.02 0.04 0.06 0.08
CL
CD
UR1T Drag Polar
Cruise CL = 0.48 Cruise CD = 0.027
K = 0.053 e = 0.77
63
would be at its minimum allowing the aircraft to takeoff and land in the shortest distance
possible. The maximum lift coefficient is 2.7 and occurs at an angle of attack of 20°. The wing
will be mounted at a 4° angle in order to reduce the tip-back angle required to takeoff. To cruise,
the lift coefficient needs to be at 0.48 which occurs at about 4° angle of attack, which is the angle
the wing is mounted at allowing the aircraft to cruise level.
Figure 39: Lift Curve Slope of UR1T.
0
0.5
1
1.5
2
2.5
3
0 5 10 15 20 25
CL
α (degrees)
Lift Curve
Plain Wing
CLmax = 2.7
Cruise CL = 0.48 Cruise α = 4°
64
16.0 Performance
16.1 Flight Envelope
The operational envelope is a description of the aircraft’s speed limits at different
altitudes. It is defined as the level flight at a specific gross weight, and is divided into three
boundary sections: minimum speed/ stall speed, minimum rate of climb/absolute ceiling, and
maximum speed/ thrust equals drag.
Figure 40: UR1T Operational Flight Envelope.
Table 11: UR1T Operational Limits.
Condition Pressure Altitude, ft. Vtrue
, knots*
Stall speed limit Sea level to 30,000 185
Absolute ceiling R/C
limit
35,500 344
Maximum speed/
Thrust=Drag limit
34,200 to sea level 433
65
To determine the first boundary, the stall speed must be determined at altitudes ranging
from sea level to 30,000 feet. The second boundary required knowing the altitude where absolute
ceiling occurs. Selected Mach numbers of 0.6 and 0.75, along with dynamic pressure at altitudes
30,000 and 35,000 feet were used to calculate the appropriate rates of climb. Further parameters
used in finding these climbs were aerodynamic data such as cruise lift coefficient at different
Mach numbers, thrust to weight ratios, drag to weight ratios, climb to cruise thrust ratios, and
cruise to sea level thrust ratios, at those particular altitudes.
The third boundary states that at maximum speed, thrust is equal to drag. Similar to the
second boundary, the parameters of interest are thrust and drag. At altitudes of 30,000 feet down
to sea level, several Mach numbers were used to locate where thrust curve intersects the drag
curve, thus yielding a specific airspeed where thrust equals drag.
16.2 Takeoff
Takeoff contributed to the design of both the wing and the engines. In order to determine
takeoff distance, takeoff was broken down into four different sections; Ground roll, rotation,
transition, and clear distance. The equations to determine these distances were taken from
Nicolai and Carichner’s Fundamentals of Aircraft and Airship Design textbook. Naturally the
ground distance is the majority of the distance since the aircraft is accelerating from zero
velocity to the takeoff velocity during this segment. The ground roll distance was calculated
using the following equation:
66
Where velocity is the takeoff velocity and acceleration is the average acceleration, this
average is the acceleration at 70.7% of the takeoff velocity (average dynamic pressure). The
above integral can be simplified to the following:
Where VTO and acceleration can be found using these equations:
The UR1T’s CLmax of 2.7 was constrained within the landing requirements and is applied
to the above equation for takeoff velocity.
67
The rotation distance was assumed to occur over a 2 second time interval. Because
rotation occurs immediately after reaching takeoff velocity, the rotation distance is simply take
off velocity multiplied by the 2 second time interval.
The transition period is the time after location where the aircraft is maneuvering from 0
degrees climb to some designated climb angle. This maneuver occurs in a circular fashion,
therefore some radius R can be calculated. Radius is dependent on the takeoff velocity as well as
the load factor that the designer sets during takeoff.
Geometry can then be done in order to determine the runway distance and height traveled during
this radial maneuver.
The UR1T aircraft was designed so that the 50 ft. tall barrier that military requires to be
cleared by the end of the runway is cleared during the transition period. This means that the
distance to clear the obstacle is zero feet.
Sg = 4217
Sr = 555
Str = 1254
Scl = 0.0
Stot = 6027
68
A balanced field length calculation was done in order to determine takeoff distance in the
event of an engine failure mid takeoff. This was done by solving for the distance it takes to
continue after engine velocity V1, as well as the distance it takes to brake. When these two
distances are equal, V1 can be solved for, and therefore the balanced field length. The sampled
balance field length was found at the 10,000ft +10C condition and the total length to takeoff was
7,750ft, falling well within the 9,000ft requirement.
69
16.3 Landing
Landing distance was solved for using a landing weight of 610,000 lbs. This weight is
the total takeoff weight minus half of the fuel. In the case of an emergency landing, a fuel dump
system will be included to pump out the fuel in order to land with this maximum landing weight.
According to FAR 25.195, the actual landing distance the aircraft is capable of must be divided
by 0.6 as a safety factor. This requirement made landing significantly more difficult and affected
CLmax as well as maximum landing weight. To determine total landing distance, landing was
broken down into 3 subcategories, approach distance, free roll distance, and braking distance.
Approach distance is simply a function of approach angle as well as touch down velocity,
where touchdown velocity is 1.15 times the stall velocity. Using an approach angle of 3 degrees,
the approach distance was found to be 960 ft.
The free roll distance is a three second period after the plane lands where it is simply rolling
without any braking; therefore this distance can be found using the following:
𝑆𝑓𝑟 = 3 ∗ 𝑉𝑇𝐷 = 3 ∗ 267 = 800 𝑓𝑡
The final braking distance is the distance it takes to go from the touch down velocity to
essentially zero, or taxiing speeds. The equation to solve braking distance was taken directly
from Nicolai and Carichner’s Fundamentals of Aircraft and Airship Design and is as follows:
The coefficient of friction used to solve braking distance was 0.30, as this was one of the
worst case scenarios for landing. The worst case flight conditions was the 10,000ft +10C
70
condition. After plugging in all the variables the total braking distance was found to be 3,540ft,
making the total landing distance 5,300ft. However, after the correction factor of dividing by 0.6
the braking distance of the UR1T was found to be 8,830ft.
16.4 Tactical Approach
From the V-n diagram, the maximum load factor was 3.0. Using a maximum bank angle
of 80°, the turn radius was calculated to be 1,400 ft going at 177 knots. The aircraft is able to
land from 10,000 ft in 4.8 minutes.
16.5 Payload Range
Two missions were specified in the RFP. One mission was to fly with 120,000lb. payload
for a range of 6,300nm and the second mission was to fly with 300,000lb. for an unspecified
range. Table 13 shows a summary of mission design. Since the second mission did not have a
range requirement, a range of 1,800nm was chosen for intercontinental flight. The ferry mission
shown is with no payload and full fuel. Figure 42 is the Payload vs Range plot. There are two
lines on this plot due to the maximum takeoff weight of the two missions being different.
Table 12: Mission Cruise Segments.
MTOW
(lb.)
Fuel
Weight
(lb.)
Fuel
Ratio
Range
(nm)
Cruise
Mach
Cruise
Altitude
(ft)
Mission 1
(120,000lb)
689,000 191,000 0.3 6,301 0.75 30,000
Mission 2
(300,000lb)
699,000 76,900 0.11 1,800 0.75 30,000
Ferry
(No Payload)
569,000 191,000 0.35 8,900 0.75 30,000
71
Figure 41: Payload vs. Range.
Table 13: Cruise Segment Specifications.
Cruise 1 Cruise 2
Altitude (ft) 22,000 30,000
Climb Time(min) 10.5* 6.6
Mach 0.75 0.75
Airspeed (knots) 457 442
L/D 16.3 17.8
Distance (nm) 1,500 4,700
0
50000
100000
150000
200000
250000
300000
350000
0 2000 4000 6000 8000 10000
Pay
load
(lb
s)
Range (nm)
Payload vs. Range
Mission 2
300,000 lb payload
Mission 1
120,000 lb payload
72
17.0 Cost Estimation
17.1 Acquisition Cost
To determine cost estimation the values were computed using the Nicolai and Carichner
methodology in Fundamentals of Aircraft and Airship Design, Volume 1. This textbook provided
the outline to calculate important cost figures including flyaway costs, and acquisition cost.
The acquisition cost required to complete an aircraft in the production line fully
operational and ready to fly. It should include the costs required in previous research and
development, including the tooling and labor required to produce all production units and
prototypes. To produce 120 units by 2030, the production line would begin early 2025 and the
labor rates and inflation rates were taken into account.
The average prototype cost was 2.65 billion dollars due to the airframe engineering, and
specialized tooling and equipment necessary for analysis and development. Other factors that
played contributors to the testing costs were engine costs, as extra engines other than the
standard four engines had to be purchased. As well as materials, test facilities, and flight test
operations. The total development and research cost for three prototypes would yield a 7.95
billion dollar initial evaluation phase.
73
Table 14: Total Cost of 3 Prototypes at 85% Learning Curve.
Prototype
Cost/Unit
Developmental
Engineering
Cost
Developmental
Tooling Cost
Developmental
Manufacturing
Cost
Material
and
Equipment
Cost
TOTAL
PROTOTYPE
COST
Average
Cost at
85% LC
1.25 Billion 466 Million 386 Million 26.5 Million 2.65 Billion
Average
Cost at
90% LC
1.31 Billion 488 Million 404 Million 27.8 Million 2.75 Billion
TOTAL
COST for 3
%85 LC
Prototypes
3.75 Billion 1.4 Billion 1.15Billion 79.6 Million 7.95 Billion
TOTAL
COST for 3
%90 LC
Prototypes
3.93 Billion 1.46 Billion 1.21 Billion 83.3 Million 8.25 Billion
17.2 Operation Cost
The flyaway cost of the four engine military aircraft was calculated using the same
methodology as the prototype costs, only considering the different phases of manufacturing and
production that occur during a program lifecycle. The production cost included figures for
sustaining engineering, tooling labor rates, quality control, and amongst the most important
manufacturing facilities. When programs with large production estimations are projected the
most efficient way to decrease error and increase production is to aim to shift more hand
assembly into machine operations, this is called applying a learning curve. For a 120 unit
production program there will not be enough product to achieve immense learning, however as
the number of man-hours decreases at the rate of learning, the cost efficiency totals over time. To
realistically estimate the manufacturing cost of the program an 85% learning curve was applied
to all components of production to get a flyaway cost of 105 Million per unit. The third party off
74
the-shelf products come at standard rates, and it is assumed that the product or service does not.
Figure 42: Total Program Savings When Learning Curves Applied at 85% and 90%.
17.3 Life Cycle Cost
The Program Lifecycle cost will experience an appreciable help in budgeting as each
product costs only 85% of the last to make. This sequence depreciates slowly resulting in the last
120th
unit only costing 78 $million as opposed to the 140 $million it cost to get the first unit of
the production line. The table below indicates the price ranges experienced at different learning
curves.
0.00
100,000.00
200,000.00
300,000.00
400,000.00
500,000.00
600,000.00
700,000.00
800,000.00
900,000.00
1 21 41 61 81 101
85% Production Engineering
85% Production Tooling
85% Manufacturing Labor
85% Quality Control
90% Production Engineering
90% Production Tooling
90% Manufacturing Labor
90% Quality Control
85% VS 90% LEARNING CURVE OF
MANHOURS PER UNIT
75
Table 15:Total Production Cost at 85%, 90%, and 95% Learning Curves.
Unit 95% Learning Curve
Unit Cost
90% Learning Curve
Unit Cost
85% Learning Curve
Unit Cost
10th 202 Million 169 Million 140 Million 60th 177 Million 129 Million 92 Million 120th 168 Million 116 Million 78.3 Million Average Unit
Cost 181 Million 136 Million 101 Million
Total Production
Cost
21.8 Billion
16.3 Billion
12.1 Billion
Figure 43: Life Cycle Cost at 85% Learning Curve.
0.00
20,000,000.00
40,000,000.00
60,000,000.00
80,000,000.00
100,000,000.00
120,000,000.00
1 21 41 61 81 101
85% Production
Engineering
85% Production Tooling
85% Manufacturing Labor
85% Quality Control
85% LEARNING CURVE OF
PRODUCTION COST PER UNIT
76
Table 16: Total Flyaway Cost of Leading Competitors.
Aircraft Model Fly Away Cost
Percent Cost
comparison of UR1T
to leading competitors
TAG UR1T 105 Million
Boeing 777 320 Million 32% of 777
Airbus A380 400 Million 26% of A380
ReEngine C5 181 Million 58% of ReEngine
When the TAG UR1T is compared to other aircrafts in the market today, the price is
considerably lower than anything available today. UR1T is 32% of the Boeing 777, and 26% of
the cost of the Airbus A380. If the C-5 were to be “re-engined” today the cost of labor and new
engines would cost about $181 Million, which is 58% of the cost to purchase a new UR1T
Military transport aircraft.
Conclusion
UR1T was designed to meet the minimum requirements of the new generation military
transport aircraft expected to be launched by 2030. The design started with constraint diagrams
that chose the initial values for wing loading and thrust to weight ratio. These parameters were
then finalized through an iteration process using a mission program specifically written for the
RFP missions. With these main parameters, trade studies were utilized to select the optimal
airfoil and engine designs for the aircraft. The empennage was designed through simple
calculations with considerations of the requirements of the RFP. Landing gear was placed based
on load distribution of the payload and other parameters such as tip back and turn around angle.
The fuselage was designed to fit the payload accordingly for both the M104 Assault Bridge and
the Master Pallets.
77
UR1T aircraft was a response to AIAA’s RFP for a next generation military transport
aircraft to be launched by 2030. UR1T is by the next generation military transport because of the
fact that it is simple and effective. All of the requirement have been met and nothing more. Any
additional implementation would be unnecessary and would only result in extra weight, in other
words, extra cost in the budget. The UR1T aircraft will be used throughout the nation and it will
outperform its previous predecessor not only in performance but utility as well. There is no doubt
in mind that UR1T will be the best aircraft of its time and with that we are team TAG and UR1T.
78
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Acknowledgements
T.A.G would like to acknowledge Grant Carichner for all the support and advice
throughout the design process. We would also like to thank all the judges for their professional
inputs.