final report -aircraft design
TRANSCRIPT
KENT STATE UNIVERSITY
AERN 45700/55700: AIRCRAFT DESIGN
INSTRUCTOR: D. BLAKE STRINGER, PH.D.
Spring 2015
The Flash by
Kent Aerospace, Inc.
Kayla Grass
Matthew Gazella Johathan Herman
Alexander Flock Steven Johns
Thomas Spisak Scott Konesky
Obed Asamoah Franklin Costa
Daniel Abbas Nicholas Brown
Guojie Wang Di Xu
Table of Contents
1. The Flash
1.1. Description
1.2. Summary of Key Parameters
1.3. Configuration Layout
2. Requirements Analysis
2.1. Requirements Summary
2.2. Mission Profile
2.3. Reference Design Concepts (Baselines)
3. Technical Design
3.1. Reference Design Concepts (Baselines)
3.2. Sizing Methodology
3.2.1. Design Space
3.3. Assumptions
3.3.1. Assumptions Used for Lift-to-Drag Ratio (L/D)
3.3.2. Assumptions Used for Initial Sizing
3.3.3. Assumptions Used for Thrust-to Weight Ratio (T/W)
3.3.4. Assumptions Used for Wing Loading (W/S)
3.3.5. Assumptions Used for Wing, Tail, and Fuselage Geometry
3.4. Wing and Tail Geometry
3.4.1. Airfoil Selection
3.4.2. Wing Geometry
3.4.3. Fuselage Geometry
3.4.4. Tail Geometry
3.5. Thrust-to-Weight Ratio
3.6. Introduction to Powerplant Data
3.6.1. Introduction to The Flash
3.6.2. Flash Performance
3.6.3. DGEN 380 Specifications and Performance
3.6.4. Static Takeoff Condition 0-15,000 Feet
3.6.5. Static Cruise Condition 10,000-23,000 Feet
3.6.6. DGEN 380 Specifications and Performance
3.6.7. Non-Static Cruise Condition
3.6.8. Non-Static Max Speed Condition
3.6.9. Wind Tunnel Data and L/D Curve
3.7. Wing Loading Data
3.7.1. Stall
3.7.2. Takeoff
3.7.3. Cruise
3.7.4. Discussion of the Wing Loading
3.8. Sizing Results and Design Selection
3.8.1. Sizing Variability and Optimization
3.9. Sizes and Capacities
3.9.1. Fuselage
3.9.2. Wing
3.9.3. Tail
3.9.4. Landing Gear
3.9.5. Fuel
3.9.6. Powerplant
3.10. Weight and Balance
3.11. Performance and Sub-System Designs
3.11.1. Flight Controls
3.11.2. Avionics
3.11.3. Electrical System
3.11.4. Landing Gear
3.11.5. Pressurization System
3.11.6. Fire Protection System
3.11.7. Fuel System
4. Manufacturing Plan
4.1. Manufacturing Readiness Levels
4.1.1. Defining Manufacturing Readiness
4.1.2. Manufacturing Readiness Levels
4.2. Industrial Base
4.2.1. Price Induction
4.2.2. Garmin
4.2.3. Rockwell Collins
4.2.4. Heroux-Devtek
5. Legal and Regulatory/Safety
5.1. FAA Certification Strategy
5.2. Risk Mitigation Strategy
5.3. Risk Identification
5.3.1. Risk Assessment
5.3.2. Risk Response Planning and Reevaluation
6. Program Management
6.1. Modification or New System
6.2. Unique Program Circumstances
6.3. Total Planned Production
6.4. Program Schedule
6.4.1. Basis for Delivery and Performance Period Requirements
6.4.2. Program Schedule
6.4.3. Activities Planned for Subsequent Phases
6.4.4. Criteria to Move into the Next Phase
6.5. Life Cycle Support
6.6. Program Management Staffing and Organization
7. Finance
7.1. Cost Estimate
7.2. Direct and Indirect Cost Estimates
7.3. Fuel Estimates
8. Value Proposition and Marketing Strategy
8.1. Competition Strategy
8.2. Sustainment Strategy
8.3. Sales and Distribution
9. Socio-Economic/Ethical Impacts
10. Conclusion
Appendices
References
The Flash
1. The Flash
1.1. Description
The Flash is considered to be a new class of aircraft; a light personal jet. The market for
this type of product is expanding and should yield high profits beginning in the third year of
production. The Flash will be marketed to small businesses, flight schools and the government,
to name a few. Its DGEN 380 Turbofan engines by Price Induction make this aircraft unique in
the sense of saving the consumer in fuel and maintenance costs as well as weight. The aircraft
was designed for the purpose of Price Induction creating a market for the sale of their engines.
The nominal cruising altitude is 18,000 feet PA and the aircraft is capable of carrying three
passengers in addition to the pilot. Its state of the art avionics package will attract many
customers and make the pilot’s job much easier.
1.2. Summary of Key Parameters
Wing Geometry
Performance Parameters
Basic Performance
Dimensions (L) 34' 5"
Engine Type DGEN380
Max Airspeed 250 kcas
Wing Span 37.34 ft
Static Thrust HP 580
Cruise Speed 230 kcas
Wing Chord 7.66 - 1.91 ft
Thrust at 18,000 ft 340
Service Ceiling 25,000 ft PA
Aspect Ratio 7.8
SFC 0.26
Range 800 NM
Wing Surface 178.76 ft²
MGWTO 4897 lbf
Endurance 3.16 hrs
Wing Loading 25 lb/ft²
1.3. Configuration layout
2. REQUIREMENTS ANALYSIS
2.1. Requirements Summary
Based upon current socio-economic drivers, the following requirements have been determined:
- The design will be 14 CFR Part 23 compliant.
- The design team will utilize Part 21 Certification procedures.
- The aircraft will utilize a fly-by-wire system to reduce weight.
- The DGEN 380 engine incorporates a FADEC system for reduced maintenance
costs as well as an electric starter for weight reduction.
- Multi-functional displays will be used in the cockpit for exceptional pilot
situational awareness.
- The aircraft will be capable of carrying 3 passengers in addition to the single pilot.
- The overall design will incorporate techniques to enable stable handling.
- Aircraft skin made of composites will further reduce weight.
- The aircraft will have a range of 800 nautical miles.
- The aircraft will be capable of short take-offs and landings.
2.2. Mission Profile
Above is the mission profile expected of the Flash with the associated fuel burns expected
for each leg, or mission segment. Leg 1. will include engine start, possible mission equipment
checks and take-off. Leg 2. includes the aircraft climbing to a cruising altitude of 18,000 feet PA,
however, it is capable of reaching 25,000 feet PA. Leg 3. is the cruise portion where with the
climb and descent and loiter portions will allow the aircraft to cover up to 800 nautical miles.
Leg 4. is the descent and loiter portion where the expected loiter time is 20 minutes. Finally, leg
5. is the landing, taxi and shutdown portion.
2.3. Reference Design Concepts (Baselines)
Eclipse 400 Dimensions _ Performance _ Powerplant PW610
length 29ft cruise speed 380mph max thrust 900lbs
wingspan 36ft range 1445mi bypass ratio 1.83
height 8ft 10in service ceiling 41000ft
empty weight 2480lbs
gross weight 4480lbs
Eclipse 500 Dimensions _ Performance _ Powerplant PW610 x2
length 33ft 1in cruise speed 380mph max thrust 1800lbs
wingspan 37ft 3in range 1295mi bypass ratio 1.83
height 11ft service ceiling 41000ft
empty weight 3550lbs
gross weight 5520lbs
Phenom 100 Dimensions _ Performance _ Powerplant PW617E-F x2
length 42ft 1in cruise speed 400mph max thrust 3390lbs
wingspan 40ft 4in range 1356mi bypass ratio 2.7
height 14ft 3in service ceiling 41000ft
empty weight 7132lbs
gross weight 10472lbs
Cirrus Vision SF50 Dimensions _ Performance _ Powerplant FJ33-5A
length 39ft 11in cruise speed 345mph max thrust 1000lbs
wingspan 38ft 4in range 1266mi TSFC 0.486
height 10ft 6in service ceiling 28000ft
empty weight 3700lbs gross weight 6000lbs
Diamond D-Jet Dimensions _ Performance _ Powerplant FJ33-4A
length 35ft 1in cruise speed 276mph max thrust 1900lbs
wingspan 37ft 9in range 1553mi TSFC 0.486
height 11ft 10in
service ceiling 25000ft
empty weight 3120lbs gross weight 5115lbs
3. TECHNICAL DESIGN
3.1. Reference Design Concepts (Baselines)
3.2. Sizing Methodology
We came upon the aircraft sizing for the wingspan, length, and height just by looking at
other aircraft of a similar category that have successfully flown and looking at what their
respective dimensions are. For the size of aircraft we are promoting in thi s project, a wingspan
from 37-40 feet seemed to be what all the successfully flown very-light personal jets have as
their wingspan. The length we came to was due to the inspirations mentioned eariler with an
average length of 35-40 ft being the most prevalent. Also the length was influenced by the
placement of the engines as we decided early on for the twin DGEN engines to be mounted to
the side of the rearward fuselage. The height was influenced by other aircraft of the same class
as before, with further influence by the seating arrangement. We needed to decide where the
passengers would sit and how tall an average person sitting in the type of seat we wanted would
equate to. The other sizing parameters such as weight and range were calculated by the class
individually and the chosen numbers taken from those that were deemed more accurate than
the rest.
3.2.1. Design Space
Since the aircraft was designed around the engines, we knew from the beginning what our
altitudes of operation would be. Price Induction had al ready determined the engines to be
operationally sound up to an altitude of 25,000 feet PA. Considering the power the DGEN 380
produces, a lighter jet was the only viable option.
3.3. Assumptions
Major assumptions affecting the design:
3.3.1. Assumptions Used for Lift-to-Drag Ratio (𝑳 𝑫⁄ )
𝐿
𝐷𝑚𝑎𝑥 estimation constant: 𝐾𝐿𝐷 = 15.5 for civil Jets
Wetted area ratio: 𝑆𝑤𝑒𝑡 𝑆𝑟𝑒𝑓⁄ = 4.1
Aspect Ratio: AR= 7.8 for General Aviation-twin engine
3.3.2. Assumptions Used for Initial Sizing
Range: R = 800 [𝑚𝑛𝑖]
Loiter Time-Endurance: E = 20 [𝑚𝑖𝑛]
Cruise Speed at FL180: 𝑀𝑐𝑟𝑢𝑖𝑠𝑒 = 𝑉𝑐𝑟𝑢𝑖𝑠𝑒 = 0.35 Mach
Constant in empty weight fraction equation: A= 1.51 for General aviation-twin engine
Constant in empty weight fraction equation: C= -0.10 for General aviation-twin engine
Variable swept constant: 𝐾𝑉𝑆 = 1.00 for fixed sweep
3.3.3. Assumptions Used for Thrust-to-Weight Ratio (𝑻 𝑾⁄ )
Maximum speed: 𝑀𝑚𝑎𝑥 = 1.2 𝑀𝑐𝑟𝑢𝑖𝑠𝑒
Constant in T/W statistical estimation equation: a= 0.267 for Jet Transport
Constant in T/W statistical estimation equation: C= 0.363 for Jet Transport
3.3.4. Assumptions Used for Wing Loading (𝑾 𝑺⁄ )
Take off distance: 𝑆𝑡 𝑜⁄ = 2500 [𝑓𝑡]
Take off Parameter: TOP = 120
Approach Speed: 𝑉𝐴𝑃𝐻 = 120 [𝑓𝑡]
Oswald Efficiency: e = 0.8
Zero-Left-Drag coefficient: 𝐶𝐷0 = 0.015 for Jets
3.3.5. Assumptions Used for Wing, Tail, and Fuselage Geometry
Taper Ratio of wing: 𝜆𝑤 = 0.25
Constant in Fuselage length equation: a = 0.67 for Jet transport
Constant in Fuselage length equation: c = 0.43 for Jet transport
Taper Ratio of tails: 𝜆ℎ= 𝜆𝑣= 𝜆𝑤 = 0.25
Aspect ratio of horizontal tail: 𝐴𝑅ℎ = 2 3⁄ 𝐴𝑅
Aspect ratio of vertical tail: 𝐴𝑅𝑣 = 1.5
Horizontal tail volume coefficient: 𝑐𝐻𝑇 = 0.90 for twin turboprop
Vertical tail volume coefficient: 𝑐𝑉𝑇 = 0.08 for twin turboprop
3.4. Wing and Tail Geometry
This section discusses the airfoil selection and parameters for geometry sizing of wings,
tails and fuselage.
3.4.1. Airfoil Selection
The selection of airfoil is one of the most critical phases in the conceptual design. The
characteristics of a specific airfoil will have a significant effect on the performance of wings. The
ideal selection is the airfoil which is capable of producing high lift and low drag. Airfoil selection
largely depends on the general considerations of the following factors:
- Airfoil geometry, such as camber and thickness;
- Aerodynamic characteristics, such as lift and drag characteristics;
- Stall characteristics;
- Other considerations, such as Reynolds number, structural layout, and different
components (Raymer, 2012).
A variety of airfoils have been developed by different institutions. In the selection of this
design, the consideration will only depend on the airfoils developed by NACA. Four ser ies of
airfoils developed by NACA are widely used in modern aircraft, the four-digit series, five-digit
series, the six-series airfoils, and seven-series airfoils. By comparing several airfoils from the
above factors, it is desirable to select the airfoil commensurate to the ideal one. However, there
are always some tradeoffs through the process of selecting.
3.4.2 Wing Geometry
Based on the TOGW determined at the initial sizing, the coefficient lift of the ideal airfoil
during cruise is determined with the ideal coefficient lift (𝐶𝑙𝑖𝑑𝑒𝑎𝑙) to be 0.18. The design lift of
coefficient is 1.11 which is the lift coefficient(𝐶𝑙𝑐𝑟𝑢𝑖𝑠𝑒) associated to the (𝐿 𝐷)𝑚𝑎𝑥⁄ . In addition,
other considerations should be included in the selection of tip airfoi l. The report Summary of
Airfoil Data published by the National Advisory Committee for Aeronautic (1945) states that it is
desirable for tip selection to have a high maximum lift coefficient ( 𝐶𝑙𝑚𝑎𝑥) and a large Critical
angle of attack (∝𝑠𝑡𝑎𝑙𝑙) in order to increase the stall performance (NACA, 1945). As for thickness,
the thicker the airfoil is, the more lift the airfoil will produce. Consequently, selecting the
thickest airfoil is advantageous.
Taking all above requirements into consideration, the criteria in response to the pivotal
factors for airfoil selections are listed below:
1. Maximum lift coefficient (𝐶𝑙𝑚𝑎𝑥) is the highest.
2. Critical angle of attack (∝𝑠𝑡𝑎𝑙𝑙) is the highest.
3. Coefficient of pitching moment (𝐶𝑚) is close to 0
4. Maximum lift-to-drag ratio (𝐶𝑙 𝐶𝑑⁄ 𝑚𝑎𝑥) at cruise is close to (𝐿 𝐷)𝑚𝑎𝑥⁄
5. Lift coefficient (𝐶𝑙) of maximum lift-to-drag ratio(𝐶𝑙 𝐶𝑑 𝑚𝑎𝑥⁄ ) at cruise is close to 𝐶𝑙𝑖𝑑𝑒𝑎𝑙
6. Minimum Drag coefficient (𝐶𝑑𝑚𝑖𝑛) is the lowest
7. Lift coefficient (𝐶𝑙) of minimum drag coefficient (𝐶𝑑𝑚𝑖𝑛) at cruise is close to 𝐶𝑙𝑐𝑟𝑢𝑖𝑠𝑒
8. Thickness ratio (𝑡 𝑐⁄ ) is highest
After comparing eleven airfoils listed in appendix 3-2, each airfoil is rated from the above
eight criteria. The airfoils with the highest rates are NACA 23012 and NACA 654-221. With
further considerations on the thickness for root and tip selections, the thickness of root section
is preferable to be thick to provide space for fuel and equipment (Abbott, Doenhoff & Stivers,
1945). According to Dr. Sadraey (2012) in his book Aircraft Design: A System Engineering
Approach, “As a guidance; the typical values for the airfoil maximum thickness -to-chord ratio
(𝑡 𝑐⁄ ) of majority of aircraft are about 6% to 18%.” (Sadraey, 2012). For different types of aircraft
in regard to speed, the maximum 𝑡 𝑐⁄ is between 9% to 12% for a high subsonic passenger
aircraft, and 15% to 18% for a low speed, high lift requiring aircraft (Sadraey, 2012). Therefore,
to optimize the airfoil to promote the performance of the designed aircraft, the airfoil NACA
23012 is selected for the tips of the wings with NACA 23015 for the roots. With the selection of
those two airfoil, the maximum coefficient of lift for wings ( 𝐶𝐿𝑚𝑎𝑥) is determined to be 1.55.
In addition to airfoil selection, there are other key factors regarding the aircraft
wings. Wing location, wing area, wingspan, and sweep angle have major effects on overall
aircraft performance.
For this aircraft, a low wing has been selected. While both high wing and low wing have
benefits, low wing is usually preferred for training purposes, and is also commonly found on
most jet aircraft. Low wing offers easier access to fueling. Low wing also allows easier access to
the engines for maintenance purposes, and allows the student to easier be able to monitor the
engine during flight. The easier access engines also help reduce maintenance costs in the long-
run, with shorter inspection times. Low wing also offers better visibility during turning and other
aerial maneuvers. Stowing landing gears is possible for both high wing and low wing aircraft, but
is much easier in low wing aircraft, as the structure is much more available to the gear.
To determine the total wing area required, it is necessary to use the following equation
(E.q.3.5.2-1). The calculated weight used in the equation is 4,897 pounds. The wing loading
calculation used is 27.3938. The resulting wing planform area comes out to be 178.76 square
feet.
𝑆 = 𝑊 / (𝑊 𝑆)⁄ ------------------------------E.q.3.5.2-1
A total wingspan of 37.34 feet has been calculated. In order to calculate wingspan, the
aspect ratio is assumed as7.8 for this calculation. The formula (E.q.3.5.2-2) explained is shown as
follows.
𝑏 = √𝐴 ∗ 𝑆 --------------------------------------3.5.2-1
Sweep angle is another important parameter regarding wing design. Changing the
sweep angle has many effects on performance, such as stability due to shifting the MAC of the
wing, or helping to avoid the onset of shock waves. From historical statistics (Raymer, 2012), a
sweep angle of approximately 2.0 degrees would be sufficient for the given aircraft.
3.4.3 Fuselage Geometry
The layout of a fuselage is generally dependent on the TOGW and the function of the
aircraft. The primary function of the designed aircraft is to carry passengers. Given the number
of passengers and crews, the length and diameter of the fuselage will eventually be determined.
However, since the proposed aircraft is also designed to undertake some other tasks more than
carrying passengers, other considerations should also be taken into account. From the historical
statistics, the following equation (Eq. 3.5.3-1) will be used to determine the length of the
fuselage:
𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒 = 𝑎𝑊0𝐶---------------------------------------Eq. 3.5.3-1
The TOGW has been determined. Based on the major assumptions made in section 3.4.5,
the length of the fuselage is calculated to be 25.87 [𝑓𝑡]. The maximum fuselage diameter is
determined by the ratio between fuselage length and maximum fuselage diameter, which is
referred as fineness ratio. To minimize the drag produced by the fuselage, the fineness ratio is
around 3 (Raymer, 2012). As a result, the maximum diameter of the fuselage is set to be 8.62
[𝑓𝑡].
3.4.4 Tail Geometry
The major function of the horizontal tail is to create a nose up moment to counter
the nose-down moment created by the wings. When the elevator or rudder is not deployed, the
tail is expected to produce zero or little amount of lift or moment. To achieve these two
purposes, symmetric airfoils are suitable selections. To several general aviation aircraft, the
NACA 0012 and the NACA0009 are applied for tails. Additionally, out of the consideration for
compressibility effect, the tails’ thickness should be less than the thickness of the wings
(Sadraey, 2012). Given the reasons above, the tail airfoil for the new design is chosen to be
NACA 0009.
The configuration of a tail is influenced by trimming, stability, controllability,
operational requirements, airworthiness and some other limits. To properly apply the
configuration of a tail requires professional analysis on the above factors. Most GA aircraft and
airline aircraft use conventional tail because it provides some benefits such as light weight,
efficient, and performs at regular flight conditions (Raymer, 2012). With limited budgets and
manufacturing level, the conventional tail will be employed in the designed aircraft.
The geometry of a tail is determined by its primary function. The tail geometry is directly
related to the wing geometry. Besides, the tail size is also related to the length of the fuselage
and the position of the engines. The tail arm is about 50% to 55% of the fuselage length for an
aircraft with the engines on the wings, about 45% to50% for aft-mounted engines (Raymer,
2012). With respect to the drawing of the new design, the engines of proposed aircraft are
mounted on the side of aft-fuselage. Therefore, the arm lengths of horizontal tail (𝐿𝐻𝑇) and
vertical tail (𝐿𝑉𝑇) are decided to be 50% of the fuselage length (𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒).
The method of calculating the parameters of tails is pertaining to the definition of tail
volume coefficient. The following two equations, Eq.3.5.4-1 and Eq. 3.5.4-2 define the horizontal
tail volume coefficient and the vertical tail volume coefficient respectively:
𝑐𝐻𝑇 =𝐿𝐻𝑇𝑆𝐻𝑇
𝐶𝑊𝑆𝑊----------------------------------Eq. 3.5.4-1
𝑐𝑉𝑇 =𝐿𝑉𝑇𝑆𝑉𝑇
𝑏𝑊𝑆𝑊-----------------------------------Eq.3.5.4-2
With the results of wing geometry calculations and assumptions made in section
3.4.5, the area of the horizontal tail is 59.55 [𝑓𝑡2] and the area of the vertical tail is 41.30 [𝑓𝑡2].
The method to calculate the other tail parameters, such as root chord, tip chord, span, length of
the MAC, and location of the AC is the same as the method used for wing geometry calculation.
Those parameters will be listed under section 3.11.3.
3.5. Thrust-to-Weight Ratio
The wings were primarily designed to support stabile handling and long endurance applications.
The thrust to weight ratio is calculated to be 25 lb/ft².
3.6. Introduction to Powerplant Data
Performance is one of the most sought after factors when developing a new aircraft. It does
not matter what kind of aircraft: helicopter, airplane, military, transport, or cargo. You will
always rely on the performance of the aircraft to complete the task at hand. The
mission/objective could be taking passengers from Chicago to New York, or a Military joint strike
fighter needing to take off from a carrier to drop a payload over a conflict zone in another
country. Each mission has its own set of established performance parameters that the aircraft
needs to meet in order to successfully complete the objective. When in the design phase, it is
necessary to list each mission the aircraft being designed needs to complete so you can start to
analyze what kind of performance will have to be met.
3.6.1. Introduction to The Flash
Our team of engineers and designers at Kent State had to immediately address the
performance factors for our aircraft. This is because we had to design the entire plane around a
DGEN 380 turbine engine. This turbofan engine has already been designed, developed, and has
begun testing to confirm its airworthiness. After it has been certified, it is then ready to move
onto the production phase. We are working with Price Induction to help design and develop a
light personal jet that will revolutionize the light jet industry. Price Induction has also developed
the Solutions Westt CS/BV DGEN 380 engine simulator. This piece of technology is a test engine
bench where the user can record and analyze different parameters occurring inside the engine
to gain a better understanding of its propulsion properties. This also gives the users the
possibility to design an entire airplane around this test engine bench. This gave our team at Kent
State an advantage because we were able to simulate what kind of performance parameters the
engines will be exposed to while the aircraft is completing its intended mission.
3.6.2. Flash Performance We are working together with Price Induction on this project of designing a new personal
light jet, so we already knew what kind of engines we would be using. Their engineers have
created a highly efficient engine that has a bypass ratio of 7.6, and is very lightweight coming in
at only 175 pounds. This engine utilizes a Full Authority Digital Engine Control or FADEC for its
power management. According to Price Induction, an all-electric concept has been validated,
such as an electric starter and ignition system. This is very critical because the onboard
generator is capable of producing 6 kw of power where 1.5 is needed for engine components
and 4.5 can be used for various airframe systems such as avionics, hydraulics etc. They have
created this engine to revolutionize the personal light jet market.
The turbofan engine was our biggest single limiting factor when designing the flash, as we
had to come up with an aircraft design that would perfectly fit these engine’s performance and
characteristics. When creating the charts and graphs of the performance we split it up into two
categories each with their own distinct parameters we set to cover a broad range of scenarios
that the Flash would be exposed to in a normal mission profile. For Static performance, we
simulated the engines to experience different altitudes with the corresponding standard
temperatures, however we did not measure the performances with the velocity of the airplane.
We calculated the static performance at both 100% throttle for a take-off condition, and also at
43% throttle for cruise conditions. Non-Static performances, again we simulated the engines
performance at the same altitudes, but included the velocity of the aircraft in the set of
parameters we were able to change on the test engine bench. For non-static we performed the
take-off, cruise and also max speed conditions.
There were also some discrepancies in our data research that we found. One of the
challenges we were faced with was dealing with a simulator that only recorded the data and
parameters using the metric system. Before we could start analyzing the data we collected, we
had to convert thrust, fuel consumption, specific fuel consumption and other parameters that
we collected to the English system. The biggest hurdle that came about was when we were
looking at the specific fuel consumption. After converting from kgFuel/kgThrust/hr to
lbFuel/lbThrust/hr. We then began to look at a calculated version of specific fuel consumption
using the equation Fuel consumption/Total thrust to see how this data compared. I found these
numbers to be completely different. We are still unsure of whether this is an issue with the
simulator or with the data we collected to input into that equation.
Within the past week, our team was able to complete a finished 3D model of the Flash and
was able to put it in Kent States wind tunnel. From the data we collected for drag forces, w e are
able to come up with an estimated thrust required curve that will be required for our aircraft at
different velocities and altitudes. Again, I would like to stress that this is just an estimate
because our model did not have the smoothest surface, which will add to the parasitic drag of
the 3D printed aircraft. Also because this is a scale model that has a scale of 1:58 inches it is very
hard to precisely calculate how the actual airplane will perform. There are many different
factors that go into each test for scale models and may not be the same factors or conditions
testing actual light jet aircraft. Another example of this could be that our 3D printed model has
fully covered engines that do not allow air to flow through them. This will greatly increase the
drag of the Aircraft.
3.6.3. DGEN 380 Specifications and Performance
Condition Thrust
Specific Fuel Consumption
Thrust At Take Off power (SLS, Mach: 0) 570 lbf 0.44
Thrust at Max Continuous (FL100, Mach: 0.338) 240 lbf 0.78
Thrust at Max Continuous (FL180, Mach: 0.4) 185 lbf 0.80
-Table 3.1
*These are performances for only one DGEN 380 Turbofan Engine.
Price Induction has come up with 2 standard applications for this engine listed below.
Standard Applications: 2 Seats (Single Engine) 4+1 Seats (Multi Engine)
Max Take off Weight: 1,980 lb 3,640 lb
Wing Loading: 25 lb/ft^2 25 lb/ft^2
Entire Surface area of A/c: 380 ft^2 700 ft^2
Max Cruise Airspeed: 247 mph 288 mph
Take Off Distance: 1,575 ft 1,900 ft
Fuel Onboard: 550 lb 1,050 lb
Range at Cruise (FL120) 615 Nm + 45 minutes 600 Nm + 45 min
Range at Cruise (FL220) 810 Nm + 45 minutes 800 Nm + 45 min
3.6.4. Static Takeoff Condition 0-15,000 Feet
3.6.5. Static Cruise Condition 10,000-23,000 Feet
3.6.6. DGEN 380 Specifications and Performance
3.6.7. Non-Static Cruise Condition
After analyzing the data for our cruise condition, we have come to a conclusion on why there is
a significant increase in both thrust and fuel consumption. We believe that this is because Price
Induction has designed this aircraft to be at optional performance at around 12,000-16,000 feet.
This characteristic is also prevalent in some of the other conditions the engine was exposed to.
3.6.8. Non-Static Max Speed Condition
9000
11000
13000
15000
17000
19000
21000
23000
25000
325 375 425 475
Alt
itu
de
(ft
)
Fuel Consumption (lbf/hr)
Fuel Consumption At Max Speed 100%
Fuel Consumption
(lbf/hr)
3.6.9. Wind Tunnel Data and L/D Curve
Creating a 3D printed model of the aircraft we designed gave us a much better understanding
of how our aircraft will actually perform in real life conditions. Various data was collected in
preparation to create a Lift/Drag curve more commonly known as the thrust required curve.
Further analysis was performed to calculate the coefficient of drag for the 3D model. This is just
an estimate, and may not be quite as high of a number on the real Flash after it is certified and
produced. These calculations were also performed at SLS conditions with the air density being
0.00237 slugs/ft3 . The higher Velocities created more accurate coefficients of drag, so the main
focus will be on those numbers.
This is our 3D printed model before it was sanded made smooth. Adam Zuckerman and
some members of our aircraft design team spent countless hours to perfect the surface of our
model in order to get it ready for the wind tunnel. This was done to lessen the parasitic drag
that will be produced from rough surfaces. As you will see below in table 3.4, our parasitic drag
was incredibly high. This led to a very high thrust-required needed to overcome this drag.
Wind Tunnel: Collected Drag Force Data
Velocity (fpm) Velocity (fps) Drag Force (lbs) Drag Force (grams)
600 10.0 0.0022 1
1150 19.2 0.00441 2
1350 22.5 0.00882 4
1500 25.0 0.01102 5
1600 26.7 0.01102 5
2090 34.8 0.01543 7
2400 40.0 0.01764 8
2600 43.3 0.01984 9
2800 46.6 0.02425 11
3000 50.0 0.02866 13
FD DragForce
AirDensity 0.00237(slugs/ ft3)
V Velocity V (Fps)
A PlanformArea 0.0523 ft2
CD 2 0.01102
0.00237 (26.72) 0.0523
CD 0.249
CD 2 0.01543
0.00237 (34.82) 0.0523
CD 0.205
CD 2 0.01764
0.00237 (402) 0.0523
CD 0.1779
CD 2 0.01984
0.00237 (43.32) 0.0523
CD 0.171
CD 2 0.02425
0.00237 (46.62) 0.0523
CD 0.1797
CD 2 0.02866
0.00237 (502) 0.0523
CD 0.185
The averages of these drag coefficients are what will be used when creating the thrust-required
curve for the 3D printed model. Again it is important to note that these characteristics will vary
for the actual aircraft since some estimation was involved in the process.
CD 2 FD
V 2 A
CDA 0.249 0.205 0.1779 0.171 0.1797 0.185
6
CDA 0.1946
From this average drag coefficient, we are now able to calculate the parasitic and
induced drag produced by our aircraft. This will then be used to calculate the thrust that is
required to overcome this drag in steady level flight.
Altitude
Density
(rho) S Weight
Oswald's
e
Aspect
Ratio K pi CD
MSL 0.00237 178.76
4897
lbf 0.8 7.8 0.05101108 3.1416 0.1946
From this calculated data, we can now create a thrust-required curve that our aircraft will need
to meet for steady level flight.
3.7. Wing Loading Data The wing loading, 𝑊 𝑆⁄ is the ratio of weight to the wing reference area. Certain
performances of an aircraft, as stall speed, rate of climb, takeoff and landing distance, lift
produced by wings, etc. are affected by wing loading. To determine the wing loading for
designed aircraft, the wing loading must be compared at some common conditions. The
following sections will present the discussion on the calculations of wing loading at three
different conditions, stall, takeoff, and cruise.
3.7.1. Stall
The stall speed of an aircraft is directly determined by the wing loading and maximum
lift coefficient. Stall speed is one of the major safety factors that need to be paid special
attention to in aviation. Several fatal accidents occur annually due to fail ure to maintain flying
speed. To determine the wing loading required to meet a certain stall speed, lift must equal
weight. Derived from the lift equation at stall condition ( E.q. 3.8.1-1), the wing loading
requirement can be determined.
𝑊 = 𝐿 = 𝑞𝑠𝑡𝑎𝑙𝑙𝑆𝐶𝐿𝑚𝑎𝑥=
1
2𝜌0𝑉𝑠𝑡𝑎𝑙𝑙
2 𝐶𝐿𝑚𝑎𝑥----------------E.q. 3.8.1-1
The formula for wing loading requirement for stall gives a result of 44.79 [𝑙𝑏𝑓 𝑓𝑡2]⁄ . This
calculation is also done with 𝐶𝐿𝑚𝑎𝑥 of 1.55, a stall velocity of 155.83 fps, and air density of
0.0024[𝑠𝑙𝑢𝑔 𝑓𝑡3]⁄ at sea level standard (SLS).
3.7.2. Takeoff
To determine the required wing loading to meet a given takeoff distance requirement,
the following expression (E.q.3.8.2-1) is used. In this calculation, the assumed takeoff distance is
2,500 feet. The takeoff parameter (TOP) can be found from fig 5.4 in the Raymer text, Aircraft
Design: A Conceptual Approach (Raymer, 2012).
𝑊 𝑆⁄ = (𝑇𝑂𝑃)𝜎𝐶𝐿𝑇/𝑂(𝑇 𝑊⁄ )𝑇/𝑂 --------------------E.q.3.8.2-1
The wing loading requirement for takeoff comes out to be 29.96[𝑙𝑏𝑓 𝑓𝑡2]⁄ . The calculated
𝐶𝐿𝑇/𝑂 is 1.281. Other variables used in the equation are the TOP which is assumed to be 120,
density ratio of 1, and (𝑇 𝑊⁄ )𝑇/𝑂 of 0.1949.
3.7.3. Cruise
Determining a wing loading for cruise is utmost important. The cruise condition is
typically the most designed around factor on an aircraft. Choosing a wing loading factor that
directly suits the cruise condition for a maximum range is problematic. The wing loading factor
for a maximum range is much higher than the wing loading factor required for stall and other
characteristics. It would be unsafe to fly with such a small wing, hence where understanding the
importance of trade-offs comes into play. To calculate the wing loading for maximum range, the
following equation (E.q.3.8.3-1) is to be used.
𝑊 𝑆⁄ = 𝑞 √𝜋𝐴𝑒𝐶𝐷0/3 ---------------------------E.q. 3.8.3-1
The dynamic press at the cruise condition is determine by the air density at cruise
altitude (FLl80) and the cruise speed. For jet aircraft, the Oswald efficiency (e) and the zero-lift
drag coefficient (𝐶𝐷0) are statistically assumed to be 0.8 and 0.015 respectively. After taking all
the variants into the above formula, the wing loading at the cruise condition is calculated to be
27.39[𝑙𝑏𝑓 𝑓𝑡2]⁄ .
3.7.4. Discussion of the Wing Loading
To determine the end wing loading requirement, all different flight operations must be
considered, such as stall, landing, takeoff, and cruise. To pick the exact wing loading that will be
used for the design process, the lowest calculated from all of the flight conditions is to be used.
After the comparing the results of the above calculation, the required wing loading is
27.39[𝑙𝑏𝑓 𝑓𝑡2]⁄ . Selecting the lowest wing loading implies that the aircraft has enough lift being
produced by the wing, for the given weight.
3.8. Sizing Results and Design Selection
3.8.1. Sizing Variability and Optimization Vary the wing loading by plus/minus 20% and the aspect ratio by plus/minus 20% to
determine the optimum combination using the carpet plot method of Chap.19.
3.9. Sizes and Capacities
3.9.1. Fuselage
Fuselage Length: 25.86 [𝑓𝑡]
Fuselage maximum diameter: 8.62 [𝑓𝑡]
3.9.2. Wing
Wingspan: 37.34 [𝑓𝑡] Root chord: 7.66 [𝑓𝑡]
Surface area: 178.76 [𝑓𝑡2] Root chord thickness ratio: 15%
Wetted area: 732.92 [𝑓𝑡2] Tip chord: 1.91 [𝑓𝑡]
Taper ratio: 0.25 Tip chord thickness ratio: 12%
LE Sweep angle: 2 [degree] MAC length: 5.36 [𝑓𝑡]
Aspect ratio: 7.8 MAC location: 7.47 [𝑓𝑡]
3.9.3. Tail
- Horizontal Tail
Root chord: 5.41[𝑓𝑡] Aspect ratio: 5.2
Tip chord: 1.35 [𝑓𝑡] Arm length: 12.93 [𝑓𝑡]
Span: 17.60 [𝑓𝑡] Taper Ratio: 2.5
Area: 59.55 [𝑓𝑡2]
- Vertical Tail
Root chord: 8.39 [𝑓𝑡] Area: 41.29 [𝑓𝑡2]
Tip chord: 2.10 [𝑓𝑡] Aspect ratio: 1.5
Span: 7.87 [𝑓𝑡] Arm length: 12.93 [𝑓𝑡]
Taper ratio 0.25
3.9.4. Landing Gear
The landing gear are designed to have a total added height of 16 inches to the aircraft.
With a tricycle type gear configuration, each strut will have a single Type III (low pressure)
wheel. More detail will be given in a later discussion.
3.9.5. Fuel
The fuel system is capable of holding 986 lbf of Jet-A fuel or 147 gallons. More detail will
be given in a later discussion.
3.9.6. Power Plant
The Flash features two DGEN 380 engines mounted aft of the wings. Each engine weighs
175 lbf and is 4 feet, 5 inches in length. More detail will be given in a later discussion.
3.10. Weight and Balance
The weight and balance of an aircraft is of upmost importance. It is important from the
very beginning of the flight until the aircraft is back on the ground. Proper weight and balance
ensures the safety of the flight and allows ease of maneuverability. The operator of a light
aircraft such as the Flash will need to closely monitor the weight and balance throughout the
flight’s entirety as the limits can be easily exceeded and thus have detrimental effects. The
following derivations are based upon statistics and chapter 15 of the Raymer text. Note: For the
most accurate information, the aircraft must be built and weighed for a proper weight and
balance to be derived.
With the diagram above, the following table was formulated as a statistical model of the
expected weight and balance of the Flash.
Weight lbs
Loc ft
Moment ft-lbs
Weight
lbs Loc ft
Moment ft-lbs
Structures 2296.9 35037.5 Equipment 466.84 7460.72
Wing 1109 14.5 16080.5 Flight controls 105 15 1575
Horizontal tail 130.5 30 3915 Hydraulics 15 0
Vertical tail 78 30.5 2379 Pneumatics 7 11 77
Fuselage 787 13 10231 Electrical 180 23 4140 Main landing gear 130 16 2080 Avionics 45 4 180 Nose landing gear 40 6 240 Furnishings 80 12 960
Firewall 22.4 5 112 Air conditioning 39.84 8 318.72
Empty weight allowance 10 21 210
Propulsion 566 13135 Total weight empty 3329.74 16.69 55633.22
Engines - installed 448 25.5 11424 Fuel system/tanks 118 14.5 1711 Useful load 1568 20918
Crew 150 9 1350
Fuel - usable 946 14.5 13717
Fuel - trapped 10 14.5 145
Oil 12 25.5 306
Passengers 450 12 5400
Takeoff gross weight 4897.74 15.63 76551.22
3.11. Performance and Sub-System Designs The design of each of the subsystems are in accordance with §23 of Federal Aviation
Regulations for Aviation Maintenance Technicians (FAR AMT). This is only a general overview of
the equipment and system operation of the major subsystems and is not all inclusive. That is to
say this section does not outline all of the requirements laid forth in the FAR AMT.
3.11.1. Flight Controls
Flight controls are essential to control the aircraft in all aspects of flight. The flight
controls modify the aerodynamic surface of the wing and in turn change the lift and drag
produced by the surface it affects. The result rotates the aircraft around one of three, or a
combination of the three, axes to change the flight path of the aircraft. The three axes and the
corresponding flight controls are the lateral axis, longitudinal axis and vertical axis
corresponding to the pitch, roll and yaw controls respectively. Pitch control utilizes the
horizontal stabilizer (horizontal tail surface), roll control utilizes the ailerons (control surface
hinged on the trailing edge of the wings), and yaw control utilizes the vertical stabilizer (vertical
control surface attached to the trailing edge of the vertical tail).
The Flash will feature the most current and pilot friendly control surfaces that will create
the ease and comfort of flight. The Flash will be using differential pressure ailerons, where one
aileron goes up than the other aileron will deflect down. This will create a more significant
change in lift and drag and a stronger roll over the longitudinal axis. The ailerons will also be
slotted, in order to add additional energy to the boundary layer. At the trailing edge of the
ailerons will be trim tabs, also controlled by the command of the pilot in the cockpit. These are
small movable portions of the control surface that alter the camber of the wing so that the
change in the deflection will hold the aircraft in an aerodynamic force. There will be balance
tabs located on the same control surface as the trim tabs; the ailerons. This tab aids in the
movement of this surface. The flaps wil l be a fowler flap. The fowler flap is a type of slotted
flap. This flap will change the camber of the wing and also increases the wing area by sliding the
flap backwards on tracks. The Flash will use a fully movable horizontal stabilizer with anti -servo
tabs. The anti-servo tab is installed on the trailing edge of the control surface and assists in
holding the control surface in its new position rather than helping it move. This will decrease
the need for additional actuators. There will be a conventional hinged rudder located on the
trailing edge of the vertical stabilizer. These control surfaces are operated through physical
commands from the cockpit controls made by the pilot. These commands are relayed to the
flight control surface through several different possible means including mechanical, hydraulic
and fly-by-wire.
The Flash will be using a fly-by-wire system. Fly-by-wire, in terms of our application is
an electrical primary flight control system (EPFCS) which is defined by the United State s Air
Force as, “a flight control system mechanization wherein the pilot’s control commands are
transmitted to the moment or force producer only via electrical
wires.” The key features that are associated with fly-by-wire systems are the replacement of
heavy hydraulic systems with electrical wires and computer assisted auto stabilizers. The fly -by-
wire system reduces the fuel costs, increase passenger capacity, has lower maintenance costs,
improves flight efficiency and reduces the fatigue of the pilot. All flight and trim controls go
through a transducer, where it will roll or pitch, and physical commands become encoded. The
encoded information is sent to the control computer which deciphers the information and sends
out commands to the surface actuators. The control computer also contains aircraft motion
sensors, which is also taken into account and makes adjustments so the pilot does not have to
conduct extra work to fulfill the flight path he wants. The command output from the control
computer also goes through servo valves attached to the actuator. The EPFCS fly-by-wire
system can contain multiple layers of redundancy to increase its reliability, without the tradeoff
weight, cost and maintenance.
The fly-by-wire system will contain built in test equipment, which will quickly detect and
isolate failures in the system. This places an added layer of safety and also decreases the
amount of maintenance man hours by directing the mechanic to the source of the failure. North
American Rockwell Corporation estimates that a fly-by-wire system will decrease the downtime
of an aircraft by at least 3% and a reduction in control system and maintenance man hours can
be reduced by as much as 80% or more. With a fly-by-wire system, the control computer can
improve aircraft handling qualities by adjusting the stick feel to the pilots preference for all flight
conditions. This is due to the control and stability augmentation. Control augmentation is
referenced to the removal of the mechanical link from the pilot to the series of servos for the
fly-by-wire operations. Pilot input is sent to the command model and data from the aircraft
motion sensors are then compiled and sent to the servo amplifier. The series servo receives
information from both the servo amplifier and the friction hysteresis before sending the
compiled instructions to the surface actuator. The stability augmentation is often referred to
the damper. Aircraft motion sensors send data to a servo amplifier which is then sent to a series
servo. The pilot input is sent to the friction hysteresis and is also sent to the series servo before
all the information is sent to the surface actuator. Both forms of augmentation are through the
same fly-by-wire system. In the fly-by-wire system, the series servo is protected by a valve from
the aircraft motion sensors and the surface actuator is also protected by a valve so as to not
cause structural damage to the mechanism.
The primary flight control computer is the PFCC-4100 from RockWell Collins, Inc. It will
be located in the nose of the aircraft. It offers very high integrity command outputs to the
actuation system. It will have stable augmentation and envelope protection. It coordinates
flight control system maintenance to ensure the quality of the flight control system. This flight
control computer is multi-channeled to evaluate many inputs, which allows one computer to
operate many redundant systems. This will ensure the safety of the flight control system. The
actuators will be from Moog. The primary flight controls will be customized fly-by-wire that will
come in dual redundant designs. All redundant systems will be ran through the primary flight
control computer. All information that goes through the flight control computer originates from
the commands generated in the cockpit and from the cockpit control. These controls include
the control column, side stick and pedals.
The design of a fly-by-wire cockpit layout is determined on the intended use of the
aircraft. Depending on purpose the customer may choose the control column or side stick as
part of their options and the panel design depends on their needs. The control column is
suggested for training purposes while the side stick is recommended for experienced pilots. The
control column style is recommended for the trainer design. There are control actuators that
provide realistic feedback to the pilot so they may experience maneuvers. This configuration
does increase weight and require more room than the side stick configuration. However, for
trainer purposes it is recommended to teach new pilots feedback from traditional control
columns. The side stick style is recommended for experienced pilots for its ease of use and
reduced weight. The side stick is adapted for emergency situations and prevents the pilot from
performing maneuvers outside of the aircraft’s capabilities. Due to its position and small size,
the side stick is more comfortable and provides an unobstructed view of the control panel.
Cockpit panels are arranged depending on use and need of the pilot. The location of the
controls takes into account each systems’ importance, the frequency of a system’s operation,
the ease that the controls can be reached and the shape of the control.
3.11.2. Avionics
The Avionics package in The Flash is primarily supplied by the Garmin G1000. The G1000 is
the premiere glass cockpit and the industry leader in crew resource management and reliable
operation. When selecting the avionics package, Garmin presents the most robust out of box
solution, which includes built in redundancies, an efficient user interface, and the modular
ability to include a wide variety of auxiliary units, called Line Replaceable Units (LRUs). In
addition to these built in redundancies, our avionics package will also include a third layer of
redundancy, in a small trio of traditional mechanical gauges operating on a completely separate
subsystem.
Image: GDU 1040
The G1000 displays all of its information through its two 10.4 inch displays, one of which
is a Primary Flight Display (PFD), and the other is designated the Multi-Function Display (MFD).
They are both the same unit, GDU 1040, and they are designated based upon their physical
location in the cockpit. The PFD is the unit directly in front of the pilot position, and the MFD is
in front of the copilot position. They are both fully customizable as to what can be displayed on
either screen, and are redundant to each other. In the event of failure all pertinent informat ion
be displayed on any single screen. These displays present information such as the artificial
horizon, heading, VOR heading, wind speed, and engine outputs, among other information.
Image: GRS 77 AHRS
The first level of information processing is the GRS 77 Atitude, Heading, and Reference
Unit (AHRS). It receives input from the GMU 44 Magnetometer, as well as it's built in tilt sensors,
accelerometers, and rate sensors. The AHRS is the primary source for aircraft attitude and flight
characteristics information.
Image: GDC 740 ADC
The primary computing center of the system is the GDC 740 Air Data
Computer(ADC). The ADC receives the input from the Pitot-Static probe, the GTP 59 OAT Probe,
as well as the AHRS and Integrated Avionics Units. The ADC determines seven primary
parameters: Total Air Temperature, Pressure Altitude, Indicated Airspeed, Calibrated Airspeed,
Vertical Speed(Rate of Climb), and Mach.
The GEA 71 Engine/Airframe Unit(EAU) provides the system with connection to the
engines FADEC and airframe sensors, such as fire sensors. It communicates with the system by
RS-485 digital communication lines.
Image: GIA 630 IAU
The heart of the system are the two GIA 630 Integrated Avionics Units(IAUs). The IAUs
provide the displays with their functionality, as well as contain the GPS receiver, NAV radio
receiver, and communications transceiver. The two units provide each other with redundancy. If
one unit fails, the other senses this failure, and all tasks are handled by the functioning unit until
the failed unit is replaced. This redundancy covers all functions with the exception of GPS, which
requires both units to achieve the required accuracy.
The IAUs communicate with the other components through a variety of
communication lines. The displays communicate to each other, as well as the IAUs through
standard Ethernet. The IAUs communicate to all of the other components through ARINC 429, as
well as RS-232. The use of two communication styles provides automatic error correction
through comparison.
The primary audio interface for the G1000 is the GMA 13470 Audio Panel. It controls
all audio controls, including intercom radios, NAV radio, communications radios, and optional
XM radio. It is mounted between the displays, and communicates only with the IAUs across RS-
232.
The GTX 330 Transponder is a Mode-S transponder which provides modes A, C, and S
ATC communication. It is controlled by the IAUs through the display.
There are many optional LRUs which can be incorporated into the G1000, which can add
a variety of features and provide more robust functionality based upon the customer's needs.
Additional features include XM Radio, Weather systems, and any number of different displays.
The basic configuration of the G1000 provides the necessary functionality to fully equip an
aircraft for flight, and represents the cutting edge of modern glass cockpit.
Images: Artificial Horizon, Indicated Air Speed, Pressure Altitude
In addition to the G1000, the aircraft will also be equipped with a completely separate set
of traditional units. These units include Artificial Horizon, Indicated Air Speed, and Altitude
displays. They are included in our avionics package to create a third layer of redundancy, which
will protect the aircraft in the event of a full G1000 system failure.
3.11.3. Electrical System
The electrical system on The Flash is a parallel-type bus system operating at 400 Hz 115
Volts Alternating Current, in one of three phases, and 28 Volts Direct Current. The parallel bus
arrangement allows for immediate failure detection and prevents the aircraft from losing full
power in the event of incidents such as single engine failure. It consists of two parallel
subsystems, named Left and Right, which is named based upon the engine powering the
subsystem, as viewed from the pilot's perspective. The Left and Right systems are connected at
two points, the first is the Main AC Bus, and the second is the Main DC Bus. The system also
allows for the input of a ground power cart or truck, which is a separate AC subsystem.
In the event of an engine or component failure in any one AC subsystem, the
functioning subsystem can automatically and quickly transfer its power into a secondary
subsystem. The secondary subsystems contain many of the same components as the main
subsystem, but feed completely separate buses. These buses, while separate, can power all of
the essential components to the aircraft.
The engine's primary electrical output control is it's incorporated Full Authority
Digital Engine Control (FADEC) software. The first component which physically begins generating
electricity is the Integrated Drive Generator (IDG), mounted onto the engine, and controlled by
the FADEC. The engine produces electricity based on its current operating conditions, and the
it's the responsibility of the FADEC to control the IDG in order to prevent potentially harmful
situations from concurring. The IDG utilizes the Generator Control Current Transformer (GCCT),
communicating through a load controller, to condition the AC power into the acceptable range.
The GCCT transforms the Alternating Current (AC) power directly from the generator
into 400 Hz 115 Volts AC (VAC), and into the appropriate phase with the rest of the system. The
initial phase is determined by the first power source operating during that run cycle, and all
other AC systems conform to it for the duration of the run cycle.
The Generator Breaker (GB) is the first breaker in line from the generator. Its primary
function is to prevent common failures from affecting any components which are still operating.
The GB also provides the connection between the Main AC Bus, the IDG, and the secondary AC
bus.
The Bus Tie Breaker (BTB) provides the cutoff point for the main AC sub system, and
closes the circuit in case of failure. The BTB, GB, and IDG all communicate with the GCU, which
monitors the current flow in the subsystem and can determine failure.
The Differential Protection Current Transformer (DPCT) is a comparative transformer
which uses the method of differential protection to monitor the current flowing through both
the primary and secondary subsystems, as well as the IDG output. While the GCCT monitors the
system as a whole, and communicates with the Load Controller and GCU, the DPCT monitors the
miscellaneous feeder wires for shorted and open conditions. The purpose of the DPCT is not
necessarily to transform the engine output, but it regulates the components drawing from the
system.
Completely separate from either of the IDG powered lines is the Ground Power
subsystem. This line allows for the use of a cart or truck to supply the aircraft with power, when
it is available. There are many advantages for utilizing the ground power availability, primarily
the ability to start the aircraft's engines without the need to draw energy from the battery. The
subsystem contains two DPCTs for line protection, and an Engine Pressure Ratio(EPR) sensor,
which is provided the current engine output. This EPR sensor closes the subsystem when the
engines have reached a self-sustaining operation, which protects the ground cart from damage
due to substantial back loads.
The Direct Current(DC) subsystem is again divided into three further subsystems,
based upon where their power is rectified from. The primary DC subsystem begins at the
Transformer/Rectifier Unit(T/R U) main, which draws directly from the Main AC Bus. Similarly,
the T/R U Left and Right draw their power from the corresponding AC Buses. Al l power
conditioning for the DC subsystems is handled by the T/R Us, which convert the AC power into
24 V DC power.
Surge protection in the DC subsystems is primarily provided by semiconductor
diodes. These diodes prevent DC power from traveling in a reverse path of the intended flow, as
well as protect the system from excessive voltages. Fuses
From the T/R U power is routed into the Essential DC Bus, which powers components
such as the Battery Bus, and the avionics. It's placement before the Main DC Bus creates a
possibility for protection from failure which may occur in the Main DC Bus. This arrangement is
an attempt to limit the effect of, say, light failure, from affecting flight essential components
such as the avionics.
The Battery Bus provides power from the battery into the Essential DC Bus when
there is need for it. The most common need for battery power is when there is no ground cart in
place for the system before and during engine start. Battery power is also a final backup for
flight in the event of dual engine failure, and according to FAA regulations the battery must be
able to provide power for the aircraft for 30 minutes. The aircraft will contain two 12 V Lithium-
Ion batteries connected in series to fulfill this requirement.
While Lithium-Ion (Li-Ion) batteries are still relatively new in aviation, their use has
become more accepted. Companies such as EaglePicher have fully developed FAA registered Li -
Ion batteries and chargers, which provide a substantial increase over traditional Lead-Acid
batteries in the areas of weight, cycle variability, and discharge duration.
The battery bus also houses the battery charger, which receives power from the AC
subsystems. While physically separate from the DC subsystem, the battery charger converts AC
power into DC power in order to charge the batteries, which then powers the DC system as
described in the conditions above. While the batteries store 24 V of power, the battery charger
provides 28 V of power to charge the batteries.
Image: Electrical System Diagram
3.11.4. Landing Gear
The Flash utilizes a tricycle type landing gear system with a single wheel per strut. This
type of gear will allow the cockpit to remain at a level attitude during taxi and takeoff as well as
allow the pilot good visibility and controllability. The landing gear is capable of withstanding up
to 90% of its max takeoff gross weight in the event of an emergency landing needing to be
performed shortly after takeoff. The landing gear will be retractable to reduce the effects of
drag and allow a smoother, faster flight. The landing gear utilizes a 12 VDC bi -directional
electro-hydraulic power pack and pump to place the gear into the desired position. The landing
gear will be produced and assembled by Heroux-Devtek Incorporation and shipped to us for
final installation onto the aircraft.
The wheel and tire selections are based upon tables, charts and equations listed in
chapter 11 of the Raymer text. The landing gear will weigh 130 lbf. The nose gear will have a 7 ̊
forward displacement to counter act any tendency for the gear to retract upon a hard landing.
Both the nose and main landing gear will utilize the same size tires and wheels. The tire will be
type III, low pressure, and can support a maximum speed if 120 mph, or 104 knots. The tires are
capable of supporting 4400 lbf, 90% of the takeoff weight. The area footprint of the tire is 90
in². The tires will have an overall diameter of 25.65 inches and a width of 8.7 inches. The tires
will be capable of holding a maximum of 55 psi. During landings, a centering cam will ensure the
nose gear is in the straight ahead position since it is the only gear capable of swiveling.
Additionally, a drag strut and side brace link will be utilized on the main gear for safety concerns
in the event of a high crab landing. Air-oleo type shock absorbers are utilized on each strut and
can be serviced via an air valve at the top of each strut. The use of an air-oleo versus a spring-
oleo allows for conservation of weight while still cushioning landings and taxiing over rough
surfaces.
When the pilot selects the landing gear to move into either the extended or position, an
electric motor is energized and rotates a cam plate that opens the landing gear stowage doors,
positions the gear, and closes the doors. Once the gear is in the selected position, a microswitch
breaks the circuit to the motor and causes the appropriate gear indication to be displayed on
the multi-functional displays. The gear will retract to stow in a fuselage-podded configuration.
For the purpose of enhanced safety, a landing-gear-position indicator system is utilized. Squat
switches allow the system to determine when the aircraft is on the ground, disallowing the gear
to be retracted accidently. A warning horn will sound when the throttle is reduced below 100
knots and the landing gear is not in the down position. In the event of a complete electrical
failure, a backup CO2 accumulator will use its charge to place the landing gear in the landing
configuration, referred to as an emergency gear up release valve blow down system.
The brakes are a single disk type and are operated via a brake-by-wire system. When
the pilot presses the brakes, an electrical signal is sent from the brake pedal transducers and the
Garmin 1000 system to actuate electrical brake actuators. This system utilizes no hydraulic
fluid, allowing for weight conservation. The brake actuators provide braking power to either
one or all wheels, at the pilot’s discretion, via pressure applied to the individual foot pedals.
Disks are rigidly bolted to the wheel and a brake housing is attached. The pistons in the brake
housing have linings on them which must be replaced when worn below tolerances, much like
the brakes of a car.
3.11.5. Pressurization System
Environmental system is typically including air conditioning system and pressurization
system. They are working together to create a comfortable atmosphere for passengers and
crews in the cabin.
For personal light jet and/or very light jet, cabin pressure differential is generally up to
6.7±0.1 psi. The preset pressure differential value is 6.8psi. This allows a sea level cabin
altitude up to 12,000 feet. And our maximum cruise altitude is 25,000 feet, so the cabin
altitude would be 5,000 feet.
The basic components include an avionics linked digital controller and two outflow
valves mounted in the aft pressure bulkhead. The MFD displays all pressurization
parameters and the PFDs provide pilot interface for entry of landing field elevation. In thi s
design, no bleeding air system is applied instead of conventional bleeding air system
coordinated with pneumatic system. Firstly, cabin air will be vented directly from the
outside through dedicated inlets on each side of the plane's belly and will not pass through
the engines. And then electrically driven compressor compresses the ram low-density air.
After that it is transported via ducts to the air conditioning packs. Within the A/C unit, the
desired temperature is achieved by regulating the adjustable speed motor compressors at
the required pressure without significant energy waste. And the regulated air distributes
through outlets in the cockpit and overhead vents in the cabin, respectively. The system
may be operated anytime in flight, or on the ground when ground power is connected or
either engine is running. A fresh air vent with a blower and a check valve is located beneath
the nose baggage compartment to provide outside air to the cockpit whenever the cabin is
not pressurized.
This approach is significantly more efficient than the traditional bleed system because it
avoids excessive energy extraction from engines with the associated energy waste by
precoolers and modulating valves. That results in significant improvements in engine fuel
consumption.
3.11.6. Fire Protection System
Fire is one of the most dangerous threats to an aircraft. Fire protection systems is very
important for every aircraft. It is installed in an aircraft to detect and protect against an
outbreak of fire. For the fire zones for our aircraft “The Flash” it will be divided into three
sections. This include the engine section, the nose compartment and the main cabin. For the
engine section, it will be divided into two zones namely zone A and zone B. Zone A is going to
cover the core section of the engine and it is also going to be provided with fire detection and
extinguishing. Zone B will cover the exhaust pipe and pylon section. One extinguisher is going to
be paced on each engine with Halon 3301 and one is going to be placed in the cockpit.
For the air cooled radial engines, the power section and all portions of the exhaust system must
be isolated from the engine accessory compartment by a diaphragm that meets the firewall
requirements of part 23.1191. The design of the fire protection for this aircraft will be in
compliance with the requirement which include:
(a) Each engine, power units and all other combustion equipment will be
isolated from the aircraft by firewalls.
(b) The firewall will be constructed so that no hazardous quantity of liquids,
gas, or flame can pass from the compartment created by the firewall.
(c) Each opening in the firewall will be sealed with close fitting, fireproof
grommets, bushing, or firewall fittings. Our firewall will be made up of composite
material. The firewall for this aircraft will be protected against corrosion and also will
be a fireproof and this is going to protect it from any danger of fire and as a result,
passengers and crew doesn’t get electrocuted when they inside of the cabin. For
example the material that will be used which requires firewall materials and fittings
must resist flame penetration for at least 15 minutes. For the design for this aircraft,
the following material will be used
1. Stainless steel sheet, 0.015 inch thick
2. Mild steel sheet (coated with aluminum or otherwise
protected against corrosion) 0.018 inch thick
3. Monel metal, 0.018 inch thick
4. Steel or cooper base alloy firewall fittings
5. Titanium sheet, 0.016 inch thick
All aircraft have an extinguishing system. The kind of extinguisher that is going to be used on this
aircraft will be the class B which is more effective with flammable liquids and with chemicals
that include monoammonium phosphate and sodium bicarbonate. The next type of extinguisher
that will be used is class C which is suitable for fire in electrical equipment with chemicals that
include monoammonium phosphate and sodium bicarbonate.
Image: Fire Suppression Bottles for Engines
3.11.7. Fuel System
All powered aircraft require fuel on board to operate the engines throughout the phases of
flight. A fuel system consists of storage tanks, pumps, valves, filters, fuel lines, monitoring
devices, and metering devices. Each system must provide an uninterrupted flow of contaminant
free fuel regardless of the aircraft’s attitude or flight condition. Varying fuel loads and shifts in
weight during maneuvers must not negatively affect control of the aircraft in flight. In general,
fuel systems must be constructed and arranged to ensure fuel flow at a rate and pressure
established for proper engine functioning under each likely operating condition. It also must be
designed and arranged to prevent the ignition of fuel vapor within the system by direct lightning
strikes.
For multiengine aircraft, each fuel system must be arranged so that, in at least one system
configuration, the failure of any one component does not result in the loss of power of more than
one engine. If two fuel tanks interconnected to function as a single fuel tank, there must be
independent tank outlets for each engine, and each incorporating a shut-off valve. The shutoff
valves may serve as firewall shutoff valves. Lines from each tank outlet to each engine must be
completely independent of each other. The fuel tank must have at least two vents arranged to
minimize the probability of both vents becoming obstructed simultaneously. In addition, aircraft
fuel tanks must be designed to retain fuel in the event of a gear-up landing. In case of sever
emergency situations, there must be a means to allow flight crew members to rapidly shut off the
fuel to each engine individually in flight.
The Flash has two fuel tanks that can carry a combined 147 U.S Gallons of Jet A fuel,
consisting of one tank per wing. Each wing has fuel receptacle that is located above the wing root
behind a spring loaded cover flap. Each receptacle then consists of a fueling nozzle adapter and
sealing cap. From the receptacle a fuel line runs downward into each respective fuel tank. There
are two primary fuel pumps in each tank located at opposite sides of the respective tank to allow
for continuous supply of fuel to the engine during maneuvers when the aircraft’s attitude is not
level. These two fuel pumps flow into one singular fuel line at a T-joint with one-way valves
preventing backflow returning to the fuel tank. Secondary or backup fuel pumps are located
adjacent to the primary fuel pumps; one secondary pump for each primary fuel pump. They use
most of the same fuel lines as their adjacent primary pump. The secondary pumps are on standby
until activated by the pilot, or if fuel pressure drops below a certain amount, they will be
automatically switched on. A collector box in the wing root keeps the electrical pumps inlets
submerged. To prevent pump cavitation, a pump and flaps valves ensure enough fuel in the
collector box at all times.
A single fuel line connects each tank with a crossfeed valve located along the centerline of
the fuselage. An air valve located above the fuel pump allows air to be vented outside for priming
the crossfeed line at engine startup, and allows for air to be pumped into the crossfeed line at
engine shutdown to prevent unwanted expansion of fuel during times of engine inactivity. There
is also a fuel vent system with vent tanks located at the wing tips which prevent damage to the
wings due to excessive buildup of positive or negative pressures inside the fuel tanks and to
provide ram air pressure within the tanks. For fuel indication within the cockpit, four fuel sensors
are installed inside each tank, and are equally spaced across the full length of the tank. measure
fuel levels at each sensor’s location and send the information to a computer that constantly
calculates the overall fuel level of the tank. For manual measurement, there are direct measuring
sticks located on the wings.
4. MANUFACTURING PLAN
4.1 Manufacturing Readiness Levels
Matters of manufacturing readiness and producibility are as important to the successful
development of a system as those of readiness and capabilities of the technologies intended for
the system. Their importance has long been recognized in the Department of Defense (DoD)
acquisition, and are reflected in current DoD acquisition policies. For an aerospace company, it is
very beneficial to follow the DoD standards and practices.
4.1.1 Defining Manufacturing Readiness
According to the DoD, Manufacturing Readiness is the ability to harness the
manufacturing, production, quality assurance, and industrial functions to achieve an operational
capability that satisfies mission needs in the quantity and quality needed by the aircraft to
perform as it is designed to at the "best value." Best value refers to increased performance as
well as reduced cost for developing, producing, acquiring, and ope rating systems throughout
their life cycle. Timeliness also is important. Our aircraft, "The Flash" must maintain a
technological advantage over our competitor's aircraft. This requires efficient development and
acquisition cycles for advancing technologies.
Manufacturing Readiness begins before, and continues during the development of an
aircraft's systems, and continues even after a system has been in the field for a number of years.
The ability to transition technology smoothly and efficiently from development, production, and
deployment into the field is a critical enabler for evolutionary acquisition.
Manufacturing Readiness Levels (MRLs) are designed to be measures used to assess the
maturity of a given technology from a manufacturing prospective. The purpose of MRLs are to
provide decision makers with a common understanding of the relative maturity, and attendant
risks associated with manufacturing technologies, products, and processes being considered to
meet DoD requirements.
4.1.2 Manufacturing Readiness Levels
There are ten MRLs that are correlated to nine Technology Readiness Levels (TRLs) in use.
The ten MRLs are described in detail below. In regards to production of the aircraft, at MRL 8,
low rate initial production can begin. At MRL 9, there is the capabili ty to go into full rate
production. By MRL 10, full rate production is demonstrated and lean practices for efficient
production are in place.
According to the National Aeronautics and Space Administration (NASA), TRLs are a type
of measurement system used to assess the maturity level of a particular technology. Each
technology project is evaluated against the parameters for each technology level and is then
assigned a TRL rating based on the projects progress. There are a total of nine technology
readiness levels. TRL 1 is the lowest and TRL 9 is the highest.
MRL 1: Basic Manufacturing Implications Identified
This is the lowest level of manufacturing readiness. The focus is to address
manufacturing shortfalls and opportunities needed to achieve program objectives. Basic
research begins in the form of studies.
MRL 2: Manufacturing Concepts Identified
This level is characterized by describing the application of new manufacturing concepts.
Applied research translates basic research into solutions for broadly defined needs. Typically this
level of readiness in the Science and Technology (S&T) environment includes identification,
paper studies and analysis of material and process approaches. An understanding of
manufacturing feasibility and risk is emerging.
MRL 3: Manufacturing Proof of Concept Developed
This level begins the validation of the manufacturing concepts through analytical or
laboratory experiments. This level of readiness is typical of technologies in categories of
research, development, and materials processes have been characterized for manufacturability
and availability, but further evaluation and demonstration is required. Experimental hardware
models have been developed in a laboratory environment that may possess limited
functionality.
MRL 4: Capability to produce the technology in a laboratory environment
In this level, technologies should have matured to at least TRL 4. This level indicates that
the technologies are ready for the development phase of acquisition. At this point, required
investments, such as manufacturing technology development, have been identified. Processes
to ensure manufacturability, producibility, and quality are in place and are sufficient to produce
technology demonstrators. Manufacturing risks have been identified for building prototypes and
mitigation plans are in place. Target cost objectives have been established and manufacturing
cost drivers have been identified. Producibility assessments of design concepts have been
completed. Key design performance parameters have been identified as well as any special
tooling, facilities, material handling and skills required.
MRL 5: Capability to produce prototype components in a production relevant environment
This level of maturity is typical of the mid-point in the development phase of acquisition.
Technologies should have matured to at least TRL 5. The industrial base has been assessed to
identify potential manufacturing sources. A manufacturing strategy has been refined and
integrated with the risk management plan. Identification of enabling critical technologies and
components is complete. Prototype materials, tooling and test equipment, as well as personnel
skills have been demonstrated on components in a production relevant environment, but many
manufacturing processes and procedures are still in development. Manufacturing technology
development efforts have been initiated or are ongoing. Producibility assessments of key
technologies and components are ongoing. A cost model has been constructed to assess
projected manufacturing cost.
MRL 6: Capability to produce a prototype system or subsystem in a production relevant
environment
For MRL 6, technologies should have matured to at least TRL 6. It is normally seen as the
level of manufacturing readiness that denotes completion of S&T development and acceptance
into a preliminary system design. An initial manufacturing approach has been developed. The
majority of manufacturing processes have been defined and characterized, but there are still
significant engineering and/or design changes in the system itself. However, preliminary design
of critical components has been completed and producibility assessments of key technologies
are complete. Prototype materials, tooling and test equipment, as well as personnel skills have
been demonstrated on systems and/or subsystems in a production relevant environment. A cost
analysis has been performed to assess projected manufacturing cost versus target cost
objectives and the program has in place appropriate risk reduction to achieve cost requirements
or establish a new baseline. This analysis should include design trades. Producibility
considerations have shaped system development plans. Long-lead and key supply chain
elements have been identified.
MRL 7: Capability to produce systems, subsystems, or components in a production
representative environment
At this level, technologies should be on a path to achieve TRL 7. System detailed design
activity is underway. Material specifications have been approved and materials are available to
meet the planned pilot line build schedule. Manufacturing processes and procedures have been
demonstrated in a production representative environment. Detailed producibility trade studies
and risk assessments are underway. The cost model has been updated with de tailed designs,
rolled up to system level, and tracked against allocated targets. Unit cost reduction efforts have
been prioritized and are underway. The supply chain and supplier quality assurance have been
assessed and long-lead procurement plans are in place. Production tooling and test equipment
design and development have been initiated.
MRL 8: Pilot line capability demonstrated; Ready to begin Low Rate Initial Production
This level is entering into Low Rate Initial Production (LRIP) of the aircraft. Technologies
should have matured to at least TRL 7. Detailed system design is essentially complete and
sufficiently stable to enter low rate production. All materials are available to meet the planned
low rate production schedule. Manufacturing and quality processes and procedures have been
proven in a pilot line environment and are under control and ready for low rate production.
Known producibility risks pose no significant challenges for low rate production. The engineering
cost model is driven by detailed design and has been validated with actual data.
MRL 9: Low rate production demonstrated; Capability in place to begin Full Rate Production
At this level, the system, component or item has been previously produced, is in
production, or has successfully achieved low rate initial production. Technologies should have
matured to TRL 9. This level of readiness is normally associated with readiness for entry into Full
Rate Production (FRP). All systems engineering design requirements should have been met such
that there are minimal system changes. Major system design features are stable and have been
proven in test and evaluation. Materials are available to meet planned rate production
schedules. Manufacturing process capability in a low rate production environment is at an
appropriate quality level to meet design key characteristic tolerances. Production risk
monitoring is ongoing. LRIP cost targets have been met, and learning curves have been analyzed
with actual data. The cost model has been developed for FRP environment and reflects the
impact of continuous improvement.
MRL 10: Full Rate Production demonstrated and lean production practices in place
This is the highest level of production readiness. Technologies should have matured to
TRL 9. Engineering design changes are minimal, and generally limited to quality and cost
improvements. Systems, components or items are in full rate production and meet all
engineering, performance, quality and reliability requirements. Manufacturing process
capability is at the appropriate quality level. All materials, tooling, inspection and test
equipment, facilities and manpower are in place and have met full rate production
requirements. Rate production unit costs meet goals, and funding is sufficient for production at
required rates. Lean practices are well established and continuous process improvements are
ongoing.
Although the MRLs are numbered, the numbers themselves are unimportant. The
numbers represent a non-linear ordinal scale that identifies what maturity should be as a
function of where a program is in the acquisition life cycle.
Level Definition DoD MRL Description
1 Basic Manufacturing Implications Identified
Basic research expands scientific principles that may have manufacturing implications. The focus is on a high level assessment of manufacturing opportunities. The research is unfettered.
2 Manufacturing Concepts Identified
This level is characterized by describing the application of new manufacturing concepts. Applied research translates basic research into solutions for broadly defined military needs.
3 Manufacturing Proof of Concept Developed
This level begins the validation of the manufacturing concepts through analytical or laboratory experiments. Experimental hardware models have been developed in a laboratory environment that may possess limited functionality.
4
Capability to produce the technology in a laboratory environment
This level of readiness acts as an exit criterion for the MSA Phase approaching a Milestone Decision. Technologies should have matured to at least TRL 4. This level indicates that the technologies are ready for the Technology Development Phase of acquisition. Producibility assessments of design concepts have been completed. Key design performance parameters have been identified as well as any special tooling, facilities, material handling and skills required.
5
Capability to produce prototype components in a production relevant environment
Mfg. strategy refined and integrated with Risk Management Plan. Identification of enabling/critical technologies and components is complete. Prototype materials, tooling and test equipment, as well as personnel skills have been demonstrated on components in a production relevant environment, but many manufacturing processes and procedures are still in development.
6
Capability to produce a prototype system or subsystem in a production relevant environment
This MRL is associated with readiness for a Milestone B decision to initiate an acquisition program by entering into the EMD Phase of acquisition. Technologies should have matured to at least TRL 6. The majority of manufacturing processes have been defined and characterized, but there are still significant engineering and/or design changes in the system itself.
8
Pilot line capability demonstrated; Ready to begin Low Rate Initial Production
The system, component or item has been previously produced, is in production, or has successfully achieved low rate initial production. Technologies should have matured to TRL 9. This level of readiness is normally associated with readiness for entry into Full Rate Production (FRP). All systems engineering/design requirements should have been met such that there are minimal system changes. Major system design features are stable and have been proven in test and evaluation.
9
Low rate production demonstrated; Capability in place to begin Full Rate Production
The system, component or item has been previously produced, is in production, or has successfully achieved low rate initial production. Technologies should have matured to TRL 9. This level of readiness is normally associated with readiness for entry into Full Rate Production (FRP). All systems engineering/design requirements should have been met such that there are minimal system changes.
10
Full Rate Production demonstrated and lean production practices in place
Technologies should have matured to TRL 9. This level of manufacturing is normally associated with the Production or Sustainment phases of the acquisition life cycle. Engineering/design changes are few and generally limited to quality and cost improvements. System, components or items are in full rate production and meet all engineering, performance, quality and reliability requirements. Manufacturing process capabil ity is at the appropriate quality level.
According to NASA, the following are the Technology Readiness Levels mentioned above are
displayed below.
4.2 Industrial Base
At this stage, our Manufacturing Readiness Level is currently at Level 2, and then will be
proceeding into Level 3.
2 Manufacturing
Concepts Identified
This level is characterized by describing the application of new
manufacturing concepts. Applied research translates basic
research into solutions for broadly defined military needs.
3 Manufacturing Proof of
Concept Developed
This level begins the validation of the manufacturing concepts
through analytical or laboratory experiments. Experimental
hardware models have been developed in a laboratory
environment that may possess limited functionality.
Setup will take approximately 12-24 months.
Suppliers:
5 Price Induction (2 DGEN 380 engines)
6 Garmin (Avionics)
7 Rockwell (Flight Controls)
8 Héroux-Devtek (Landing Gear)
4.2.1 Price Induction
Price induction is one of the few companies to have developed a modern aeronautical
gas turbine in the past decade. Its state-of-the-art product is the DGEN 380 engine, the world’s
smallest turbofan intended for 4-5 seat Personal Light Jets. This high bypass ratio geared
turbofan was designed from a blank sheet to allow for the advent of a new class of aircrafts on
the general aviation market. After fifteen years of development, the engine is recognized as a
technical success and has now to enter the certification and industrialization phase.
Price Induction’s adventure began in 1997, when Bernard Etcheparre, a French
entrepreneur, decided to launch the DGEN program to contribute to the innovation in the
general aviation market. Launched as a venture project, with a team of young engineers, the
program quickly gained the support of French aerospace laboratories, major French
aeronautical companies and institutional investment funds.
On October 31st 2006, the first DGEN 380 engine was successfully ignited with the test
benches. In 2011, the DGEN 380 completed its first 150-hour endurance block test. From 2010
onwards, in order to leverage its know-how, the company diversified its activities: the first
WESTT SOLUTIONS test bench was installed in 2011 and the first R&T project was signed in
2012. Since then, the DGEN program has undergone more than 2,000 cycles, 1,500 hours of
operations and two successful 150-hour endurance block tests. DGEN engines are regularly
produced for both the development of the program and the WESTT SOLUTIONS product family.
DGEN 380 Engine Cutaway 1
4.2.2. Garmin
Garmin's mission is to be an enduring company by creating superior products for
automotive, aviation, marine, outdoor, and sports that are an essential part of our customers’
lives. Garmin's vision is to be the global leader in every market, and the products will be sought
after for their compelling design, superior quality, and best value. The foundation of Garmin's
culture is honesty, integrity, and respect for associates, customers, and business partners. These
3 words "Build to Last" describe the products, company, culture and the future. As a leading
worldwide provider of navigation, Garmin is committed to making superior products for
automotive, aviation, marine, outdoor and fitness markets that are an essential part of our
customers’ lives.
Garmin's vertical integration business model keeps all design, manufacturing, marketing
and warehouse processes in-house, giving them more control over timelines, quality and
service. Their user-friendly products are not only sought after for their compelling design,
superior quality and best value, but they also have innovative features that enhance the lives of
the customers.
DGEN 380 Flow Visualization 1
Garmin G1000®
The Standard in Glass Flight Deck Capability
Certified on a broad range of aircraft models
Integrates virtually all avionics
See clearly even in IFR conditions with SVT™
GFC 700 digital autopilot integration
The G1000 is an all-glass avionics suite designed for OEM or custom retrofit installation on a
range of business aircraft. It is a seamlessly integrated package that makes flight information
easier to scan and process. Its revolutionary design brings new levels of situational awareness,
simplicity and safety to the cockpit.
The G1000 puts a wealth of flight-critical data at a pilot's fingertips. Its glass flight deck
presents flight instrumentation, navigation, weather, terrain, traffic and engine data on large-
format, high-resolution displays. It features a flexible design, G1000 adapts to a broad range of
aircraft models. It can be configured as a 2-display or 3-display system, with a choice of 10" or
12" flat-panel LCDs interchangeable for use as either a primary flight display (PFD) or multi-
function display (MFD). An optional 15" screen is also available for even larger format MFD
configurations.
The G1000 replaces traditional mechanical gyroscopic flight instruments with super-
reliable GRS77 Attitude and Heading Reference System (AHRS). AHRS provides accurate, digital
output and referencing of your aircraft position, rate, vector and acceleration data. It’s even
able to restart and properly reference itself while your aircraft is moving. The G1000 also
includes the GFC 700, the first entirely new autopilot designed and certified for the 21st century.
The GFC 700 is capable of using all data available to G1000 to navigate, including the ability to
maintain airspeed references and optimize performance over the entire airspeed envelope.
4.2.3. Rockwell Collins
Rockwell Collins is a pioneer in the design, production and support of innovative
solutions for their customers in aerospace and defense. Rockwell's expertise in flight-deck
avionics, cabin electronics, mission communications, information management and simulation
and training is strengthened by their global service and support network spanning 150
countries. Working together, their global team of nearly 20,000 employees shares a vision to
create the most trusted source of communication and aviation electronics solutions.
Rockwell's aviation electronics systems and products are installed in the flight decks of
nearly every air transport aircraft in the world. Their communication systems transmit nearly 70
percent of U.S. and allied military airborne communications. Whether developing new
technology to enable network-centric operations for the military, delivering integrated
electronic solutions for new commercial aircraft or providing a level of service and support that
increases reliability and lowers operational costs for our customers throughout the world,
deliver on their commitments.
Rockwell Collins is a leading provider of flight control and navigation solutions for
commercial, military and Unmanned Aircraft Systems (UAS). Their flight control systems
expertise includes autopilot, actuation, fly-by-wire, pilot controls, and engine controllers. The
flight control products exemplify our capabilities in systems engineering, precision machining,
fabrication, and assembly of close-tolerance flight critical parts to meet design and certification
requirements. Regardless of a system’s complexity, their flight controls ensure the stability and
safety of flight operation.
Fly-by-wire systems reduce weight, improve reliability, and increase aircraft fuel
efficiency. Rockwell's fly-by-wire systems help create a familiar environment for pilots by
combining computer software and hardware to emulate the look and feel of mechanical pilot
control systems. Movements of the column, wheel, and pedals are converted to electronic
signals and transmitted electronically by wires to the control surfaces.
4.2.4. Heroux-Devtek
Héroux-Devtek Inc. is a Canadian company specializing in the design, development,
manufacture, integration, testing and repair and overhaul of landing gear and actuation systems
and components for the Aerospace market. The Corporation is the third largest landing gear
company worldwide, supplying both the commercial and military sectors of the Aerospace
market. The Corporation also manufactures hydraulic systems, fluid filtration systems, electronic
enclosures, heat exchangers and cabinets for suppliers of airborne radar, electro-optic systems
and aircraft controls. The Corporation’s emphasis on Research & Development, its systems
integration accomplishments, and its engineering prowess are increasingly making Héroux-
Devtek a preferred partner for the design, qualification and manufacture of comple te landing
gear systems
5. LEGAL and REGULATORY / SAFETY
5.1. FAA Certification Strategy
This section will give a very brief overview of the aircraft and component certification
process. By no means is this to be utilized as the sole direction for the process, but a generality
for the purpose of understanding the process.
In general, there are several phases according to the FAA for the entire aircraft approval
process. The first phase is to develop the conceptual design. The conceptual design will consist
of the overall generalities of the aircraft; no specifics.
Next, the requirements need to be identified. The product definition, identification of
associated risks and a mutual commitment to move forward with those identified by both the
FAA and the applicant are completed in this phase. Many meetings take place during this phase
and a preliminary certification board meeting is held. This is where the proposed schedule for
the entire certification process is made.
Next, the aircraft will need to be designed in accordance with proper compliances.
Specific project planning is done and a Project Specific Certification Plan (PSCP) is made. This is
the FAA’s specific compliances for what type of aircraft it is. For our purposes, we designed in
accordance with CFR §23, Airworthiness Standards for Normal, Utility, Acrobatic and Commuter
Category Airplanes. This subchapter of the Federal Aviation Regulations for Aviation
Maintenance Technicians defines requirements of each subsystem and what kind of testing they
must undergo. Other parts may be specific to larger subcomponents, however. An example
would be §33 talks about the airworthiness standards of aircraft engines.
The implementation process is where you begin to see results. The applicant must
demonstrate their compliance with the FAR AMT subsections for the particular systems, show
compliance and comformance to the previously identified requirements, and have a final
certification board meeting. This is when the aircraft will be inspected and safety analysis will
be performed.
The final phase is post-certification. This phase primarily deals with processes to ensure
continued airworthiness standards are met. This includes certificate management for the
remainder of the product’s life cycle.
To begin the process, the applicant will turn in FAA Form 8110-12 to the nearest
Certification Office, which is located in Chicago, IL. Initially, this form will be filled out
requesting a type certification.
The following is a general outline of the entire process:
• Within 2 weeks after application:
• Acknowledgement of application issued
• FAA Certrification Project Notification (CPN) issued
• Within 1 month after application:
• Project team identified (FAA and Applicant)
• Preliminary Type Certification Board Meeting (PTCBM) scheduled
• Within 1-3 months after PTCBM:
• Proposed Certification Basis G-1 issue paper prepared and processing begins (stage
1)
• PSCP drafted
• Within 4-6 months after PTCBM:
• Final Certification Basis G-1 issue paper closed
• PSCP agreed and signed, including the mutually agreed project schedule
• Within 6-9 months after PTCBM:
• All issue papers closed
• One month prior to scheduled TC/STC/Production Approval issuance:
Compliance documentation submittals should be scheduled over the course of a
project to be completed by this point in time. More than on month may be needed
in some cases, especially when submittals are not FAA Designee approved or
recommended for approval
The following is identification of the key players throughout the process and their primary roles:
FAA and Applicant’s Management – Provides a commitment to the Partnership for a
Safety Plan as well as provides leadership and resources
FAA and Applicant’s Project Managers – Jointly orchestrates the project and applies the
Partnership for Safety Plan agreements
FAA Standards Staff Project Officers – Provides a timely, standardized policy and
guidance
FAA and Applicant’s Engineers and Designees – Apply regulations and policy to find
compliance including the determination of the adequacy of type design and
substantiation data
FAA and Applicant’s Inspectors and Designees – Determines conformity and
airworthiness
FAA and Applicant’s Flight Test Pilots and Designees – Conducts FAA flight tests
FAA Chief Scientific and Technical Advisors (CSTA) – Provides expert advice and technical
assistance
FAA Aircraft Evaluation Group – Evaluates conformance to operations and maintenance
requirements
Below is an example of the PSCP process:
5.2. Risk Mitigation Strategy
Risk is a function of likelihood multiplied by the severity. As long as one of the variables
in the function is rated to be high, the project will be considered to be risky. The risk
management approach includes four phases, risk identification, risk assessment, risk response
planning, and risk evaluation.
5.3. Risk Identification
In the first stage of risk management, it is crucial to pinpoint the risks and focus on
the risks that are highly likely to cause the project to fail. Risks can be found internally and
externally. The internal risks include market risk, assumption risks, and technical risks. The
project, like designing a new aircraft involves a series of high risks in the market and technical
aspects.
With the DGEN 380 engine, the new design is classified as a VLJ or PLJ. The market
for these types of aircraft is not completely exploited. With the freshness of the market, the
definition of the market remains ambiguous. In addition, when the project is delegated by the
Price Induction, the requirements from the customer need to be fully defined and include the
details to the most extent. Failure to define the market or the customer’s requirements clearly
could result in the risks of misleading the direction of the project. With the market research
made by the project team, there are three models currently on the market sold by three
different companies, but using the engines produced by the same company Pratt & Whitney
(Pratt & Whitney, n. d. ). However, a variety of aircraft in the same class are either under flying
test or in the phrase of undergoing development. Being unable to keep track of the newly
introduced products and modifying the new design could bane the competitiveness of the new
design.
Another major risks existing internally in the project is from technical aspect
including four essential features: maturity, complexity, quality, and concurrency of the project.
As a newly formed team that hasn’t dabbled in the aircraft designing for a long time, lack of
experience and knowledge could lead to the more time consuming and more expensive. With
the newly developed engine, the innovation and creativity in the project can also increase the
risks. Besides, the complexity of a project like aircraft designing can also affect the likelihood
and severity of the risks. The procedures in aircraft designed are highly related and involve
numerous interrelations. The calculations and estimations on the TOGW dictates the
calculations of the rest parameters mostly. The estimation of TOGW can be influenced by
various industrial and economic factors besides the technical factors, such as the customer’s
requirement, budget of the project, manufacturing process, and etc. Like all of the other design
projects, the end-item of aircraft design is to produce the aircraft designed by the team. In the
process of the design, the end-item cannot be completely produced or fully tested.
Consequently, the extent of testability and producibility also have effects on the risks of the
project. Last but not least, from the Gantt chart of the project, due to the time constrain on the
project, several sequential activities overlap each other and most of the activities are dependent
on the other activities.
As for external risks, the project is limited to the following factors: government
regulations, customer needs and market conditions, material or labor resources, and physical
environment. As a highly regulated industry, the project of designing a new aircraft has to be
complied with the FAA certifications and testing standards. The amount of demands and the
conditions of the market for the new product can also affect the success of the project. Lack of
materials, resources, labor forces, and terrain can also have an influence on the risks of the
project.
5.3.1. Risk Assessment
There are plenty of methods of assessing the levels of risks. Since risk is a function of
two variables, likelihood and severity. The equation (Eq. 5.2.2-1) below presents the risk
function:
Risk = Likelihood × Severity--------------------- Eq. 5.2.2-1
The method of risk matrix will be used for this project to evaluate the risks identified in section
5.2.1. The likelihood of a risk can be assessed from five levels, very unlikely, unlikely, possible,
likely, and very likely. Likewise, the severity of a risk can also be divided into five levels, low,
minor, moderate, significant, and high. The matrix below present the result of a risk considering
from both likelihood and severity:
Severity Likelihood
Low Minor Moderate Significant High
Very Unlikely Low Low Med Low Medium Medium
Unlikely Low Med Low Med Low Medium Med High
Possible Low Med Low Medium Med High Med High
Likely Low Med Low Medium Med High High
Very Likely Med Low Medium Med High High High
After evaluating each risk identified in the first phrase, risks are rated as medium high and high
will be the ones to be focused to deal with. Those risks are from the conflicts among market
conditions, product demands, and the technical concerns.
5.3.2. Risk Response Planning and Reevaluation
Five things can be arranged to manage risks, transferring, avoiding, reducing,
accepting risks, and contingency planning. Risks can be transferred by purchasing insurance and
by specifying the responsibilities and risks of each group. The groups involved in highly risky
activities should be constantly monitored by the risk management team or higher authority.
The third method of managing risk is to avoid risk. However, avoiding risks in a highly perplex
project like aircraft designing could potentially increase the complicity of the project, which
contributes to more risks. Therefore, avoiding the risk is not recommended for this project.
Instead of avoiding risks, the risks can be reduced or mitigated by reducing the
likelihood and severity of technical risks. Before putting the design into production, models and
simulations should be formed and tested to improve the performance of the design.
Additionally, the project team should always conduct a parallel development on the highly risky
tasks and assess the performance of those tasks before proceeding to the next related activities.
After the calculation on the TOGW, the project team should carefully consider a series of
conditions to refine the sizing result before using the result for further decision made on other
parameters. To critically evaluate the project before proceeding to the next one, the project
team should hire some outside consultants to assess the project. Multiple contingency plans
should also be proposed based on the scenarios brainstormed by the project team. Throughout
the whole process of the project, the risks should also be monitored. In case of new risks rising
up, the team should install the contingency plan as soon as the early symptoms of a risk show
up. Lastly, while estimating the budget of the project, the financial team should reserve a part of
budget for project delaying, cost overrun, and risk management.
6. PROGRAM MANAGEMENT
Program management is the process of managing multiple related projects at once.
Where project management is often used to describe one project, program management
involves multiple projects that are all related and working toward the same goal or result . For
the Kent Aero company, there are many advantages of using program management to manage
the separate projects that go in to completing an entire aircraft, although it can be challenging
to pull off well. Issues like governance and risk can be managed more successfully if a single
team is coordinating efforts.
Changes can be managed much more effectively as well. Completing all the related projects
within a program while staying on budget and on schedule is far more likely with good program
management than without it. The three factors that drive projects such as this are performance,
schedule, and cost.
6.1. Modification or New System
Currently, there are no active plans to modify the aircraft, add any new systems or
components. However, the option is always open as we proceed into the future. There is a
possible option in the future for entities, such as the government to purchase this aircraft, and
have it converted to suit their needs. In this case, the rear passenger seats can be removed, and
special equipment could be loaded and installed onboard.
6.2. Unique Program Circumstances
The unique circumstance for this aircraft is the fact that we are building and the design the
airframe around two Price Induction DGEN 380 turbofan engines, which are very efficient high
bypass turbofan engines, but they are still in the experimental stage, and are not fully certified
yet. The Flash must also go through a detailed FAA certification process, which was described in
the previous sections.
6.3. Total Planned Production
The Kent Aero company plans on producing an average of 2-3 aircraft per month, which
would translate to 24-36 per year. We are aiming to produce around 156 aircraft within the next
five years.
6.4. Program Schedule
Event/Process Date
Kickoff Meeting (1/16/15)
Requirements Definition (1/19/15 – 2/6/15)
Conceptual Design (1/20/15 - 2/25/15)
Preliminary Design (2/12/15 – 3/6/15)
Preliminary Design Review (3/10/15)
Detail Design (3/12/15 – 4/27/15)
Final Design Review (4/28/15)
Final Report Completion (5/5/15)
6.4.1 Basis for Delivery and Performance Period Requirements
The following are requirements for the supplier delivery of products for the aircraft, and
the resulting period requirements and guidelines. In this case, the following agreement pertains
to all suppliers of the Kent Aero Inc. company. The Buyer(s) and Supplier(s) in scenario are as
follows:
Buyer(s):
Kent Aero. Inc.
Supplier(s):
Price Induction (2 DGEN 380 engines)
Garmin (Avionics)
Rockwell (Flight Controls)
Héroux-Devtek (Landing Gear)
Complete Agreement
This Purchase Order, which includes any supplementary sheets, schedules,
exhibits, and attachments annexed hereto by Buyer (Kent Aero Inc. legal entity placing
the Purchase Order), contains the complete and entire agreement between the parties with
respect to the subject matter of this order, when accepted by acknowledgement or
commencement of performance. It supersedes any other communications, representations
or agreements whether verbal or written. The order may be accepted only on all the terms
and conditions herein stated. Additional or different terms proposed by the Supplier shall
not be applicable, unless accepted in writing by an authorized employee of the Buyer and
made a part of this order. No acceptance by Buyer of or payment for goods ordered
hereunder shall be deemed a waiver of the foregoing or an acceptance of any additional
or different terms contained in any acknowledgement, invoice or other form sent or
delivered by Supplier to Buyer. No usage or trade or course of dealing shall serve to alter
or supplement the terms and conditions herein stated.
Changes
The Buyer shall have the right to make, from time to time, changes as to packing,
testing, and destination, specifications, designs, quantity and delivery schedule of goods
covered by this order. Supplier shall promptly notify Buyer when such changes affect
price or other terms and shall request Buyer's written authorization to modify this order
accordingly. Claims for adjustments under this clause must be asserted within thirty (30)
days from the date of receipt of notification of the change(s).
Price
The price of goods covered by this order shall be as set forth on the face hereof
and shall not be subject to increase without Buyer's prior written consent.
Notwithstanding the above, the Supplier agrees that the price of such goods shall not be
less favorable than that extended to any other customer of Supplier for same or like goods
in equal quantities, and that if the price of such same or like goods is reduced prior to the
delivery of goods hereunder, the price hereunder shall be reduced correspondingly.
Unless otherwise set forth on the face hereof, the price of goods covered by this order
shall include all extra charges, including charges for packing, containers, insurance and
transportation. All taxes based upon and measured by the sales, use or manufacture and
imposed on this sale shall be shown separately on Supplier's invoice.
Delivery
Time of delivery as set forth on the face hereof is of the essence. If the Supplier
for any reason does not complete delivery of all goods covered by this order within the
time set forth on the face hereof, Buyer may, at its option, either approve the revised
delivery schedule, reduce the total quantity of goods covered by this order by the amount
of omitted shipments, reduce the price pro rata, or terminate this order by notice to
Supplier as to stated items not yet shipped or services not yet rendered and purchase
substitute items or services elsewhere and charge Supplier with any loss sustained,
without incurring any liability whatsoever for any such revision, reduction or termination.
Deliveries of goods covered by this order in advance of the time set forth on the face
hereof are prohibited without Buyer's prior written consent.
Shipping
Title to and risk of loss on all goods shipped by Supplier to Buyer hereunder shall
pass to the Buyer upon Buyer's inspection and acceptance of such goods at Buyer's plant.
All delivered goods shall be packed and shipped in accordance with instructions or
specifications of this order. In the absence of any such instructions, Supplier shall comply
with best commercial practice to ensure safe arrival at destination at the lowest
transportation cost. If in order to comply with Buyer's required delivery date it becomes
necessary for Supplier to ship by a more expensive method than specified in this order,
Supplier shall pay any increased transportation costs, unless the necessity for such
rerouting or expedited handling is due to the fault of Buyer. Numbered packing slips,
bearing the order number, must be placed in each container. Supplier must list the
packing slip number on its invoice.
Kent Aero. Inc. Requirements for Shipping and Transportation
Supplier agrees to abide by the Buyer's Shipping and Transportation
Requirements. Failure to comply with these Requirements will result in the Supplier
being responsible for the transportation costs for the above order.
Documentation Requirements for Importation
Supplier shall provide all Documentation Requirements for Importation to any
Buyer location, or shipping on our behalf. Supplier shall be responsible for any penalties
assessed by U.S. Customs on the Buyer due to non-compliant documentation.
Warranties
Supplier expressly warrants that all goods or services provided under this order
shall: (i) be wholly new and contain entirely new components and parts; (ii) be
merchantable; (iii) be free from defects in material, workmanship and packaging; (iv) be
fit and sufficient for the purpose for which they are intended; (v) conform to all
applicable specifications and appropriate standards; (vi) be equivalent in materials,
quality, fit finish, workmanship, performance and design to any samples submitted to and
approved by Buyer; and (vii) have been produced in compliance with all applicable
federal, state and local laws, orders, rules and regulations.
Supplier further warrants that it has good warrantable title to the goods, and that it
owns all patents, trademarks, trade names, trade dress, copyrights, trade secrets and other
proprietary rights (other than proprietary rights owned by Buyer) used by Supplier in
connection with the goods and services or has been properly authorized by the owner of
such proprietary rights. Supplier shall indemnify and hold Buyer harmless for all
damages arising out of any breach of these warranties. Supplier shall extend all
warranties it receives from its vendors and suppliers to Buyer, and to Buyer's customers,
and Supplier's warranties herein shall survive the delivery of goods to Buyer and any
resale of goods by Buyer. Breach of these warranties, or any other term of this order,
shall entitle Buyer to all available remedies, including those under applicable law.
Quality
Supplier shall meet all requirements in the Buyer's Supplier Quality Manual in
addition to any quality requirements detailed on the face of this order.
Inspection
All goods covered by this order shall be subject to Buyer's inspection and
acceptance at Buyer's plant or at any other place that Buyer may reasonably designate.
Buyer expressly reserves the right, without any liability hereunder or otherwise, to reject
and refuse acceptance of goods covered by this order that do not conform in all respects
to any instructions of Buyer contained on the face hereof or Buyer's specifications,
drawings, blueprints and data. Neither Buyer's payment of nor its inspection of goods
covered by this order prior to their delivery to Buyer's plant shall in any way waive
Buyer's right to make final inspection and acceptance of such goods at its plant.
Rejection
In case any goods delivered hereunder are defective in material or workmanship
or otherwise not in conformity with the drawings, specifications, samples, and/or other
descriptions or the order, such goods shall be returned to Supplier for credit or refund and
shall not be replaced or repaired by Supplier except upon written instructions from Buyer,
excepting however, those goods which Buyer and Supplier agree in writing shall be
repaired by Buyer at Supplier's expense. Any return goods shall be shipped transportation
collect (declared at full value, unless Supplier advises otherwise), and Supplier shall have
all risk of loss from and after the time of shipment. The inspection rights set forth herein
are in addition to and not in limitation of any other rights and remedies under applicable
law and the failure by Buyer to exercise its right to reject any goods shall not by
implication or otherwise cause a waiver of any such rights or remedies. Any goods
returned to Supplier for credit or refund, not repaired by Supplier, pursuant to written
instructions, shall be destroyed and not resold or disposed of to any other party or parties.
Termination
Buyer may terminate all or any part of this order at any time or times, for
convenience, by written notice to the Supplier. Supplier shall submit its termination claim
to Buyer within thirty (30) days from the effective date of termination. The provisions of
this paragraph shall not limit or affect the right of the Buyer to terminate this order for
default. Buyer shall have the right to terminate this order or any part thereof without
further cost or liability to Buyer in the event of the happening of any of the following:
filing of a voluntary petition in bankruptcy by Supplier; filing of an involuntary petition
to have Supplier declared bankrupt, if such petition is not vacated within thirty (30) days
from the date of filing; the appointment of a receiver or trustee for Supplier, if such
appointment is not vacated within thirty (30) days from the date thereof; the execution by
Supplier of an assignment of the benefit of creditors; Supplier's failure to make or delay
in making deliveries hereunder or any other failure of Supplier to perform in accordance
with this order, without excluding any other remedies available to Buyer.
In the event Buyer terminates this order, in whole or in part as provided in this
paragraph, Buyer may procure, upon such terms and in such manner as Buyer may deem
appropriate, supplies and services similar to those so terminated, and Supplier shall be
liable to Buyer for any excess costs for such similar suppliers and services. Supplier must
furnish Buyer with written notice of any cause of failure which is beyond its control and
without fault or negligence, within five (5) days of the occurrence. Upon any default or
breach of this order by Supplier, Buyer in addition to other remedies, may at its option,
require Supplier to immediately transfer to Buyer all materials, work in process,
completed goods, tooling, plans, and specifications allocable to the canceled portion of
this order.
Payment
Unless otherwise set forth on the face hereof, net invoices relating to goods
purchased hereunder shall be paid within ninety (90) days after the date of invoice or
ninety (90) days after the date of acceptance of such goods, whichever is later. Payment
for goods and/or services covered by this order will be made in the currency set forth on
the face of this order. Upon reasonable notification to Supplier, Buyer may withhold and
deduct from any part of the purchase price due under this order all or any part of the
damages including consequential damages, resulting from any breach of terms and
conditions contained herein, or any other amount which Supplier owes Buyer or any of
Buyer's associated companies.
Discounts
Cash discount period shall be computed either from date of acceptance of goods
purchased hereunder, or date of receipt of correct and proper invoices relating to such
goods, whichever date is later. Buyer shall be deemed to have paid for goods purchased
hereunder on the date on which payment is mailed to Supplier.
Intellectual Property Indemnity
Supplier warrants the goods purchased hereunder and the use of such goods by
Buyer or its customers shall not infringe or misappropriate any intellectual property
rights, including, without limitation, any copyright, trademark, trade secret, patent, or
other intellectual property right. Supplier shall defend, indemnify, and hold Buyer and its
customers harmless from any liability, or claim of liability, for such infringement or
misappropriation, including damages, costs, expense, attorney's fees and lost profits
arising from any claim or suit brought against Buyer or its customer alleging such
infringement or misappropriation, provided, however, that Supplier is notified of such
suit. In the event an injunction shall issue against Buyer in any such suit which prohibits
or limits Buyer's use of goods purchased hereunder, Supplier shall, at no cost to Buyer, at
Buyer's request, furnish Buyer with non-infringing and/or non-misappropriated
replacement goods of a similar kind and quantity or procure for Buyer the right to
continue using the original goods.
Indemnification
Supplier assumes entire responsibility and liability for any breach by Supplier of
its obligations under this Agreement and for all damage and/or injury of any kind or
nature whatsoever, (including death resulting there from) to all persons, and to all
property caused by, resulting from, arising out of or occurring in connection with
Supplier's goods sold hereunder. Except to the extent, if any, expressly prohibited by
statute, should any claims, actions and/or lawsuits for such damage, injury and/or death
be made or asserted, Supplier agrees to defend, indemnify, save and keep harmless
Buyer, its officers, agents, customers, directors, employees and affiliated companies from
and against any and all such claims, actions and/or lawsuits and further from and against
any and all loss, cost, expense, judgment, settlement liability, damage or injury, including
legal fees and disbursements, that Buyer, its officers, agents, customers, directors,
employees and affiliated companies may directly or indirectly sustain, suffer or incur as a
result thereof and the defense of any action at law which may be brought against Buyer,
its officers, agents, customers, directors, employees and affiliated companies upon or by
reason of any such claim, actions, and/or lawsuits and to pay on behalf of Buyer, its
officers, agents, directors, employees and affiliated companies upon demand, the amount
of any judgment and/or settlement that may be entered against Buyer, its officers, agents,
directors, employees and affiliated companies in any such claim, action and/or lawsuit.
Inspection of Records
Supplier agrees that all reasonable records pertaining to this order by Supplier,
shall at all reasonable times be subject to audit and inspection by any authorized
representative of the Buyer. The Supplier agrees to allow the Buyer or his representative
to inspect Supplier's facilities as required to insure order compliance.
Insurance
During the delivery of goods and during the two (2) year period following such
delivery, and during the period of any work to be performed by Supplier on Buyer's
premises, Supplier shall maintain and pay for liability insurance relating to such goods in
amounts no less than the following: (a) with respect to bodily injury liability, One Million
Dollars ($1,000,000) for each occurrence and Five Million Dollars ($5,000,000)
aggregate per policy per year; (b) with respect to property damage liability, One Million
Dollars ($1,000,000) for each occurrence and Five Million Dollars ($5,000,000)
aggregate per policy per year. The insurance shall (a) be extended to include "Vendor's
Coverage", (b) name Buyer as an additional insured and loss payee with respect to such "
Vendor's Coverage"; and (c) be written with insurance companies and contain such
provisions as shall be satisfactory to Buyer. Supplier shall furnish Buyer with certificates
of insurance confirming the existence of such insurance. All such policies shall provide
that the coverage thereunder shall not be terminated without at least thirty (30) days prior
written notice to Buyer.
Buyer's Property and other Special Tooling
Unless otherwise provided in writing or herein agreed, property of every
description, including all tooling, dies, jigs, fixtures, patterns, or other equipment and
materials furnished or made available to Supplier, or prepared by Supplier specifically in
connection with the manufacturing of goods ordered hereby, title to which is with Buyer,
and any replacement thereof, shall be and remain the property of Buyer. Property other
than materials shall not be modified without the written consent of the Buyer. Such
property shall be plainly marked or otherwise adequately identified by Supplier as
property of Buyer (by name) and shall be safely stored separately and apart from
Supplier's property. Supplier shall not use such property except for performance of work
hereunder or as authorized in writing by Buyer. Such property while in Supplier's
possession or control shall be kept in good condition, shall be held at Supplier's risk, and
shall be kept insured by Supplier, at its expense, in an amount equal to the replacement
cost with loss payable to Buyer. To the extent such property is not materially consumed
in the performance of the order, it shall be subject to inspection and removal by Buyer
and Buyer shall have the right of entry for such purposes without any additional liability
whatsoever to Supplier. As and when directed by Buyer, Supplier shall disclose the
location of such property and/or prepare it for shipment and ship freight collect on the
buyer's account.
Unless otherwise herein agreed, special tools, dies, jigs, fixtures, patterns, gauges,
molds and test equipment (hereinafter collectively referred to as "Special Tooling") to be
used in the manufacture of goods ordered hereby, furnished by and at the expense of
Supplier, shall be kept in good condition and, when necessary, shall be replaced by
Supplier, without expense to Buyer. Supplier shall at its own expense maintain such
Special Tooling and special equipment in proper working order and shall be responsible
for all loss thereof or damage thereto while in its possession and shall use the same
facilities, equipment or Special Tooling. Unless specifically provided to the contrary in
this order, Supplier warrants that the price set forth herein does not include any amount
representing rent for the use of Government-owned facilities, equipment or Special
Tooling.
Confidentiality - Information and Materials
All information and materials including, without limitation, drawings, artwork,
data, customers' names, or the like furnished by Buyer in connection with this order, shall
remain property of the Buyer and shall be used by Supplier only for work being done for
Buyer and shall be held in strict confidence by Supplier. Any knowledge or information
which the Supplier shall have disclosed or may hereafter disclose to the Buyer related to
the placing and filing of this order shall not, unless otherwise specifically agreed upon in
writing by the Buyer, be deemed to be confidential or proprietary information, and
accordingly shall be acquired free from any restrictions.
Compliance with Laws, Regulations and Supplier Code of Conduct
Supplier agrees that it will comply with all federal, state and local laws and
regulations applicable to the goods, sale and delivery of the goods or the furnishing of
any labor or services called for by the order and any provisions required thereby to be
included herein shall be deemed to be incorporated herein by reference. In addition,
Supplier agrees that it will comply in all respects of the Kent Aero Inc. Supplier Code of
Conduct. Such Code of Conduct may be amended from time to time and Supplier is
advised to review Buyer's website from time to time when it receives a new purchase
order from Buyer.
Counterfeit Parts
Supplier shall not deliver any Products to Buyer that contains any "Counterfeit
Parts" or "Suspect Parts." Such procedure may be amended from time to time. Supplier
shall indemnify and hold harmless Buyer and its officers, directors and affiliated
companies from any and all losses, damages, claims, costs and expenses for Supplier's
failure to comply with the Kent Aero. Inc. Supplier Counterfeit Material Avoidance
Procedure.
Export Laws
Supplier acknowledges that the goods and any technical data related thereto is or
may be subject to United States (U.S.), European Union (EU), or national export control
laws, regulations or the like, and agrees that it will not transfer, export or re-export the
goods or any technical data, including without limitation any documentation, or
information that incorporates, is derived from or otherwise reveals such, without
complying with all applicable U.S., EU, or national export control laws, regulations and
the like.
Manufacturer's Affidavit and Certificate of Origin
Buyer requires that Supplier complete a Manufacturer's Affidavit and
Certificate of Origin to have on file for Customs compliance matters. The
Manufacturer's Affidavit is to be filled out by Supplier's party knowledgeable of the
manufacturing of the products, or who can access the manufacturing records.
Business Continuity
Supplier acknowledges that single points of failure exist within the supply chain
and agrees to take commercially reasonable efforts to mitigate the risk of business
interruption. Efforts include, but are not limited to, the creation and implementation of a
comprehensive disaster recovery plan, periodic testing to ensure plan remains valid and
executable, and supply chain/supply base analysis and programs to eliminate exposure to
single points of failure including tooling, materials, and any other elements critical to the
manufacturing of products.
Assignment
Supplier shall not assign this order or any contract resulting here from, or any
rights hereunder, without first obtaining the written consent of Buyer. Any such
assignment without the written consent of Buyer shall, at Buyer's option, be void.
Waiver
No course of dealing between Buyer and Supplier or any delay on the part of
Buyer in exercising any rights hereunder or under any contract resulting here from shall
operate as a waiver of any of Buyer's rights, except to the extent expressly waived in
writing by Buyer.
Subcontracting
Supplier shall not subcontract any work or any goods to be supplied under this
order without the prior written approval of Buyer.
Government Subcontract
If a government contract number appears on the face of this order, Supplier agrees
to comply with all terms and conditions of that government contract.
Independent Contractor
Supplier shall perform the work necessary for performance of this contract with
Supplier's employees and agents under the control of Supplier.
Set-Off
Buyer shall be entitled at all times to set-off any amount owing at any time from
Supplier to Buyer.
Use of Buyer's Name
Supplier shall not, without first obtaining prior written consent from Buyer, in any
manner publish the fact that Supplier has furnished or contracted to furnish Buyer the
goods herein mentioned, or use the name of Buyer or any of its customers, in Supplier's
advertising or other publication. If the goods specified in the order are peculiar to Buyer's
design, either as an assembly or component part of an assembly, or if the material bears
Buyer's trademark and/or any other identifying mark, it shall not bear the trademark or
other designation of the maker or Supplier and similar material shall not be sold or
otherwise disposed of to anyone other than Buyer.
Force Majeure
Neither Buyer nor Supplier shall be liable for delay or failure of performance due
to changes in government priorities or control of materials or other necessary compliance
with changes in government regulations, or strikes, fires, accidents, acts of God, or other
causes beyond such party's control and affecting its operations. Notwithstanding the
foregoing, Buyer may terminate all or any portion of this order without liability to
Supplier if such delay or failure to perform by Supplier or on the part of Supplier extends
beyond thirty (30) days after Buyer's requested delivery date. Whenever an actual or
potential labor dispute is delaying or threatens to delay the timely performance of this
order, Supplier shall immediately give notice thereof to Buyer.
Process Control
Supplier shall make no change in material or supply chain used, construction or
fabrication techniques, test methods used without prior written approval of Buyer. Any
such changes desired by the Supplier shall be requested in writing indicating reason for
such change and including effect on cost and performance.
Severability
If any one or more of the conditions of this order shall be invalid, illegal, or
unenforceable in any respect under any applicable law, the validity, legality and
enforceability of the remaining conditions contained herein shall not be affected or
impaired in any way.
Remedies
Nothing is this order shall be claimed or deemed to limit or exclude those
remedies otherwise available to Buyer at law or in equity, and no disclaimers or
modifications or attempted disclaimers or modifications of any express or implied
warranties relating to the goods by Supplier shall be valid or effective.
*The agreement above is based off of the Kollmorgen Supplier Terms & Conditions .
6.4.2 Program Schedule
Schedule Analysis is the process of evaluating schedule results and assessing the
magnitude, impact, and significance of actual and forecast variations to the baseline and/or
current operating schedules. To date, everything has gone according to the original planned
schedule, and all current set dates have been achieved. Overall performance has been at a high
level, and all activities have been completed in an efficient and timely manner. There have been
little to no setbacks in the design process. The aircraft is in the later design phase, and will soon
proceed into the testing and production phase. During this phase, problems may arise early on,
but we are confident that our design will perform as designed, and there should be little to no
issues during that stage of testing.
6.4.3. Activities Planned for Subsequent Phases
In addition to the continued design, test, and production of the aircraft, other events
have to occur coincidently. Contracts with the supplies have to be composed and agreed upon
by both parties. Major components being installed on the aircraft have to be tested, and
validated either prior or upon arrival from the supplies. As mentioned above, Price Induction
and their DGEN 380 engines have to obtain the proper certification(s) and be considered
airworthy. §33 & §34 will have to be completed by Price Induction. The aircraft itself will have to
go through the rigorous FAA certification process. Such requirements are shown below.
FAA Certification Requirements:
o §21 Certification procedures
o §23 Airworthiness standards: Normal, utility, acrobatic and commuter category
airplanes
o §33 Airworthiness standards: Aircraft engines
o §34 Fuel venting and exhaust emission requirements for turbine engine
powered airplanes
o §36 Noise standards: Aircraft type and airworthiness certification
Shown below is the typical product certification process norms that will be occurring alongside
the major phases of aircraft design, testing and production. The product certification process
was described in detail within the Legal & Regulatory / Safety section.
6.4.4. Criteria to Move into the Next Phase
The Flash is currently near the end of the design phase, and will be proceeding into testing, and
the early stages of production. As stated previously in this report, the aircraft is currently in MRL
2, and will be proceeding into MRL 3. MRL 2 describes the application of new manufacturing
concepts. Fundamental research turns into general solutions for defined requirements.
Manufacturing producibility and overall risk is emerging. MRL 3 consists of the validation of the
concepts through analytical and lab research. Processes for manufacturability has been
characterized, but further evaluation is required. Where models have been developed for
research, and can provide some data.
The next major stage for this aircraft program is obtaining MRL 4, in which the aircraft
must be at least at TRL 4. Required investments, such as manufacturing technology
development, have been identified. Processes to ensure manufacturability, producibility, and
quality are in place and are sufficient to produce technology demonstrators. Manufacturing risks
have been identified for building prototypes and mitigation plans are in place. Target cost
objectives have been established and manufacturing cost drivers have been identified.
Producibility assessments of design concepts have been completed. Near the last phase of the
program, the aircraft will be at MRL 9-10, and will be in full production.
6.5. Life Cycle Support
A major concern for most aircraft operators is how to maintain operational capabilities
while at the same time improving availability and cost-effectiveness. One way of addressing
such a challenge is to choose Kent Aero Inc. as your support solution provider since the Flash is a
direct product of the Kent Aero Company.
With our fully integrated life cycle-based support concept we can guarantee support
solutions that will increase aircraft availability, reduce costs, and help you counter stronger
competition and greater challenges. Our commitment is long-term and includes solutions for the
entire supply chain, from factory support all the way to the airfield. With the purchase of the
Flash comes a full 2 year or 800 flight hour warranty. In addition, many components on the
aircraft require fairly low maintenance. The Price Induction DGEN 380 engines are made of parts
that can be removed per section and fairly easy, which cuts maintenance time and labor costs.
The flight controls consist of a Rockwell Collins fly-by-wire electronic control system, which
requires less maintenance than traditional hydraulic fl ight control systems. This aircraft is
designed to require less maintenance compared to similar aircraft in its class.
6.6. Program Management Staffing and Organization
Position / Area of Concentration Name
Product Lead Kayla
Requirement Analysis Kayla
Technical Design Design Team
Preliminary CAD Drawing Scott
Engine Performance Alex / Matt
Weight Performance Jon
Manufacturing Matt / Frank / Tom
Legal & Regulatory / Safety Kayla / Jasper
Program Management Matt
Finance Steven / Frank
Value Proposition Steven
Sales & Distribution Nick / Dan
Socioeconomic and Ethical Concerns / Impacts Obed / Di
7. FINANCE
7.1. Cost Estimate
The following costs of the Flash’s production are estimates based on a 5 year, 136
aircraft production line
- Engineering Costs: $124 million
-Tooling Costs: $75.5 million
-Manufacturing Costs: $185.85 million
-Quality Control Costs: $27.25 million
-Development Support Costs (Nonrecurring): $36 million
-Flight Test Costs: $11.35 million
-Manufacturing Materials Costs: $62.74 million
-Engines Costs: $108.8 million
-Avionics Costs: $6.8 million
-RDT&E+ Flyaway Total Costs: $638,440,838
7.2. Direct and Indirect Cost Estimates
-Engineering costs include airframe design and analysis, test engineering, configuration
control, and system engineering.
-Tooling costs embrace all of the preparation for production: design and fabrication of
tools and fixtures, preparation of molds and dies, programming for numerically
controlled manufacturing, and development and fabrication of production test
apparatus.
-Manufacturing Costs is the direct labor to fabricate the aircraft, including forming,
machining, fastening, subassembly fabrication, final assembly, routing, and purchased
part instillation.
-Quality control costs includes receiving inspection, production inspection, and final
inspection.
-Development support costs are the nonrecurring costs of manufacturing support
including fabrication of mockups, iron-bird subsystem simulators, structural test articles,
and various other test items used during Research, Development, Test & Evaluation
(RDT&E).
-Flight-test costs include planning, instrumentation, flight operations, data reduction,
and engineering and manufacturing support of flight testing.
-Manufacturing materials is the raw materials and purchased hardware and equipment
from which the aircraft is built.
7.3. Fuel Estimates
The Flash’s fuel consumption averages between 340-380 lbf/hr. The following fuel costs
are estimated at $4.50 per gallon as of 2015.
Fuel Costs:
-Trainer: $648,000/year
-Business Jet: $1,620,000/year
8. VALUE PROPOSITION and MARKETING STRATEGY
8.1. Competition Strategy
With the identification of the competition, we can more accurately identify our
weaknesses. We can identify the strongest market and target them. The competition will likely
change as time progresses, therefore, this will be an ongoing task for the finance and research
departments.
8.2. Sustainment Strategy
With the standing-up of the Kent Aero, Incorporation, the facility will be completely
tooled for self-sustainment. Aside from the identified products that are outsourced to be made
and assembled elsewhere, the aircraft will undergo all assembly, testing and sales at the facility.
With the ongoing research of the competition, our budget will remain fluid to ensure
the company does not fall into a deficit. This will allow longer sustainment of the company.
Additionally, over time, total costs are expected to reduce as the market expands to demand
efficiency. This will increase sales, causing a need for a larger facility.
8.3. Sales and Distribution
The Flash is designed to be affordable to a larger market and has the efficiency to justify
its cost. Potential buyers include government departments (including military application),
universities, flight schools, corporations, and independent consumers. All sales will be directed
through a chief sales consultant, whose information will be located on the corporate website
and forms of advertising that we may pursue. The marketing strategy will consist of multiple
phases to insure target exposure and rapid sales growth. We cannot target all of our potential
buyers do to such a wide spread of potential buyers, therefore we will focus our marketing on
regionally located corporations. In phase one, we will focus our marketing efforts at regionally
located corporations headquartered in Ohio, which include Kroger, Mejier, Limited Brands and
Nationwide, ect. Companies like Nationwide and Limited Brands are national brands, but
operate in regional markets, therefore by targeting them within the North Eastern Region can
lead our aircraft being sold to one company multiple times for multiple regions. Our strategy is
to expose ourselves to our target consumer. We will attend the National Business Aviation
Association Exposition and any other business aviation exposition where our target audience
will be in attendance. We will also focus our advertising in key magazines like Columbus CEO,
Chief Executive Magazine, The CEO Magazine, and others. We intend to use $500,000 in
advertising and general marketing the first year of operations. We will also utilize a discount and
endorsement strategy to begin the process of selling aircraft rapidly. For the first fifteen,
possibly more, corporations or chief executive who purchase our aircraft will receive a fifteen
percent discount if they post a public video endorsement on all forms of their social mediums. In
phase one we intend on exposing our aircraft and company to the corporate world.
We intend to sell this aircraft for its cost efficiency at the time of purchase and for its
cost efficiency over its lifetime. It’s also environmentally friendly due to its low emissions,
therefore there will be low image impact from consistent use of this aircraft. Corporations will
be able to send their executives to smaller airport for travel saving them time and money. The
executives who need to fly out for a morning meeting will be able to do so without any hassle
that comes with a larger airport and will be able to return the same day for another corporate
meeting. Our sales will be primarily from corporations initially, but our sales department will be
open to any and all sales. We do expect orders from executives who have flown in our aircraft
through their company to order one of our aircraft for their own private use. All of our aircraft
will be fitted to accompany applications for scientific equipment and transportation operations,
but it all depends on what configuration that is requested at time of purchase. The first phase of
the marketing strategy will last three to five years, depending if we keep make our target sales
goals. Phase two, will be determined during year three, but will more than likely be targeting
private flight schools and all training facilities.
9. SOCIO-ECONOMIC / ETHICAL IMPACTS
One of the main cause of environmental impact of aviation is caused by aircraft
engines which release heat, noise and gases in to atmosphere which contribute to climate
change and global dimming. Our aircraft is going to be part of the evolution of newer and
more fuel efficient engine. DGEN380 engine has been developed and improve technology
with a better fuel efficiency and reduced emissions. Low emissions from aircraft engine will
means less pollution into the atmosphere and this is what the industry and the environment
need. DGEN380 is a high by-pass ratio designed to power 4 to 5 seat light jets and can
provide up to 230 lbs. of thrust.
This engine is constructed in a sense that it use high performance materials which
allows the weight to be optimized from both a structural and a functional point of view. Air
pollution and noise go hand in hand because they both come from aircraft. Noise is
considered to be the most immediate impact of aircraft. Aircraft noise has always been a
problem for the people that work around it or the passengers that travel with it. However,
not only can DGEN380 help reduce pollution, but it can also offer low noise. The innovation
of the engine with the integration of FADEC around the engine with its high bypass ratio,
can offer a low noise and low fuel consumption. With that said, it will end up providing
efficiency for aircraft with great safety of use and comfort, low maintenance, low pollution
and a reasonable running cost. In addition, it can be used as a training aid and this can help
enhance learning experience.
However, this can be another source of bringing more businesses to the area. If
businesses can be brought from this, then this can be a way of providing more jobs in the
area of Kent, Ohio.
10. CONCLUSION
The Flash, though only a light personal jet, has many qualities that would be attractive to
our identified consumers. The overall process of designing an aircraft from the ground up is
demanding and rewarding, both. The Flash has only been designed to a readiness l evel 2,
meaning there is much more to the process yet to come until the final product is complete.
In conclusion, this overall experience has been a terrific learning experience for the entire
class. We have learned the dynamics of the collective team effort and the collaboration that
goes with. The Flash is a great concept model of what the innovative mind can accomplish with
a bit of direction. The 3-D model allowed us to determine more accurate performance
measures and to see first-hand how well our design worked out. Great futures are on the
horizon for this class.
Appendix 3-1 List of Symbols
AC Aerodynamic center
𝐴𝑅ℎ Aspect ratio of horizontal tail
𝐴𝑅𝑣 Aspect ratio of vertical tail
𝐴 Aspect ratio of wing
∝ Angle of attack
𝑏𝑤 Wing span
𝑏𝐻𝑇 Horizontal tail span
𝑏𝑉𝑇 Vertical tail span
𝐶𝐷0 Zero-Left-Drag coefficient
𝐶𝐿 Lift coefficient of wing
𝐶𝐷 Drag coefficient of wing
𝐶𝑙 Lift coefficient of airfoil
𝐶𝑑 Drag coefficient of airfoil
𝐶𝑚 Moment coefficient of airfoil
𝑐𝑤 Wing chord length
𝑐𝑡 Tip chord length
𝑐𝑟 Root chord length
𝑐𝐻𝑇 Horizontal tail volume
coefficient
𝑐𝑉𝑇 Vertical tail volume coefficient
𝐸 Endurance
𝑒 Oswald Efficiency
𝐾𝐿𝐷 Constant in 𝐿
𝐷𝑚𝑎𝑥
𝐾𝑉𝑆 Variable swept constant
𝐿 𝐷⁄ Lift-to-Drag ratio
𝐿𝐻𝑇 Horizontal tail arm
𝐿𝑉𝑇 Vertical tail arm
𝐿𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒 Fuselage length
MAC Mean aerodynamic chord
𝑀 Mach number
R Range
𝛾̅ Location of MAC
𝑆𝑡 𝑜⁄ Take off distance
𝑆𝑟𝑒𝑓 Wing Reference area
𝑆𝑤𝑒𝑡 Wetted area
𝑇 𝑊⁄ Thrust-to-Weight ratio
𝑡 𝑐⁄ Thickness-to-chord ratio
𝑡 Airfoil thickness
TOGW Takeoff gross weight
TOP Take off Parameter
𝑉𝐴𝑃𝐻 Approach Speed
𝑉𝑐𝑟𝑢𝑖𝑠𝑒 Cruise speed
𝑊 𝑆⁄ Wing loading
𝑊0 Takeoff gross weight
𝜆ℎ Taper Ratio of horizontal tail
𝜆𝑣 Taper ration of vertical tail
𝜆𝑤 Taper Ratio of wing
Appendix 3-2 Table of Airfoil Selection Comparisons
The Reynolds number for data below: 𝑅𝑒 = 3 × 106
NACA
Airfoil
Highest Highest Closer
to 0
Cruise
16.9
Cruise
1.11
Cruise
0.16
Lowest Highest
0009 1.25 13 0 112 0.8 0 0.0052 9 1
4415 1.42 13 -0.1 119 0.85 0.5 0.0075 15 0
4412 1.5 13 -0.09 125 0.85 0.4 0.006 12 0
2415 1.4 14 -0.05 122 0.82 0.3 0.0065 15 0
23012 1.6 16 -0.013 120 1 0.3 0.006 12 3
23015 1.5 15 -0.008 118 1 0.1 0.0063 15 1
631-212 1.55 14 -0.004 100 0.58 0.2 0.0045 12 1
632-015 1.4 14 0 101 0.8 0 0.005 15 1
633-218 1.3 14 -0.03 103 0.85 0.2 0.005 18 1
64-210 1.4 12 -0.042 97 0.45 0.2 0.004 10 2
654-221 1.1 16 -0.025 120 0.75 0.2 0.0048 21 3
Note: The shaded areas are the highest rates for each column.
𝑪𝒍𝒎𝒂𝒙 ∝𝒔𝒕𝒂𝒍𝒍 𝑪𝒎 𝑪𝒍 𝑪𝒅⁄
max
𝑪𝒍 of
(𝑳
𝑫)𝒎𝒂𝒙
𝑪𝒍of 𝑪𝒅𝒎𝒊𝒏
𝑪𝒅𝒎𝒊𝒏
𝒕 𝒄⁄
max
Total
rate
Appendix 5-1 FAA Certification Strategy Terms
AEG – Aircraft Evaluation Group
CFR – Code of Federal Regulations
CPI – Certification Process Improvement
CPN – Certification Project Notification
CSTA – Chief Scientific and Technical Advisor
DER – Designated Engineering Representative
FMEA – Failure Modes and Effects Analysis
FSDO – Flight Standards District Office
GAMA – General Aviation Manufacturer’s Association
JAA – Joint Airworthiness Authorities
MOPS – Minimum Operational Performance Standard
PM – Project Manager
PSCP – Project Specific Certification Plan
PSP – Partnership for Safety Plan
TC – Type Certification or Type Certificate
TSO – Technical Standard Order
REFERENCES
Abbott, I., Doenhoff, A., & Stivers, L. National Advisory Committee for Aeronautics,
(1945). Summary of airfoil data (Report NO.824)
MOOG, INC. (2013). Flight Control Actuation. 3
Nicholas, J., & Steyn, H. (2012). Project management for engineering, business and
technology (4th ed.). New York, NY: Routledge.
Raymer, D. (2012). Aircraft Design: A Conceptual Approach (5th ed.). Reston, VA: American
Institute of Aeronautics and Astronautics, Inc.
Remer, Dale. D. (1996). Aircraft Systems For Pilots. Englewood, Co: Jeppesen 3
Sanderson Training Products.
Sadraey, M. (2012). Aircraft Design: A Systems Engineering Approach. N.p.: Wiley Publications
Turbofan Engine: PW600. (n.d.). In Pratt & Whitney Canada. Retrieved March 29, 2015, from
http://www.pwc.ca/en/engines/
AcqNotes. (2015). Manufacturing Readiness Level (MRL) - AcqNotes. Retrieved from
http://acqnotes.com/acqnote/careerfields/manufacturing-readiness-levelmanufact
Airbus. [Graph Image]. Retrieved from
http://www.airbus.com/uploads/pics/Material_andLogistics_intro_section__1_.png
Cornell University Law School. (n.d.). 48 CFR 7.105 - Contents of written acquisition plans. |
LII / Legal Information Institute. Retrieved April 14, 2015, from
https://www.law.cornell.edu/cfr/text/48/7.105
DoD Manufacturing Readiness References and Links. (n.d.). Retrieved from http://www.dodmrl.com/
Garmin. (n.d.). About Us | Garmin | United States. Retrieved from http://www.garmin.com/en-
US/company/about/
Garmin. (n.d.). G1000 | Garmin. Retrieved from https://buy.garmin.com/en-US/US/in-the-air/flight-
decks/g1000-/prod6420.html
Heroux-Devtek Inc. (2010). Corporate Profile - Heroux-Devtek Designer, Developer and Manufacturer of
Aerospace & Industrial Products. Retrieved from
http://www.herouxdevtek.com/company/corporate-profile
Kollmorgen. (n.d.). Supplier Terms & Conditions | Kollmorgen. Retrieved from
http://www.kollmorgen.com/en-us/service-and-support/partners/supplier-terms-
conditions/
Majerowicz, W., & The Boeing Company. (2001, November). Schedule Analysis Techniques
[PDF]. Retrieved from http://www.evmlibrary.org/library/TP-
16%20Schedule%20Analysis%20Techniques,%20Majerowicz.pdf
National Aeronautics and Space Administration. (2014, February 3). Technology Readiness Level | NASA.
Retrieved from https://www.nasa.gov/content/technology-readiness-level/#.VTnRPSFViko
OR-AS. PM Knowledge Center. (n.d.). Schedule Risk Analysis: How to measure your baseline
schedule?s sensitivity? | PM Knowledge Center. Retrieved from
http://www.pmknowledgecenter.com/dynamic_scheduling/risk/schedule-risk-analysis-
how-measure-your-baseline-schedule%E2%80%99s-sensitivity
Pcubed. (2015). Definition: What is Program Management. Retrieved from
http://www.pcubed.com/services/glossary.program
Price Induction. (n.d.). About Us | Price Induction. Retrieved from http://www.price-
induction.com/about-us/who-we-are/
Rockwell Collins. (2015). Flight Controls. Retrieved from
https://www.rockwellcollins.com/Products_and_Systems/Controls.aspx
Rockwell Collins. (2015). Fly-by wire. Retrieved from
https://www.rockwellcollins.com/Products_and_Systems/Controls/Fly_by_wire.aspx
Rockwell Collins. (2015). Our Company. Retrieved from http://rockwellcollins.com/Our_Company.aspx
Saab. (n.d.). Aircraft Support Solutions. Retrieved from http://saab.com/commercial-
aeronautics/aircraft-support-solutions/
The FAA and Industry Guide to Product Certification, Second Edition (2004). AIA, GAMA and
the FAA Aircraft Certification Service. Retrieved from
https://www.faa.gov/aircraft/air_cert/design_approvals/media/CPI_guide_II.pdf.
WebFinance, Inc. (n.d.). What is program management? definition and meaning.
Retrieved April 29, 2015, from http://www.businessdictionary.com/definition/program-
management.html
Aircraft Pressurization Systems. (n.d.). Retrieved April 6, 2015, from
http://www.tpub.com/ase2/75.htm
http://acqnotes.com/acqnote/careerfields/manufacturing-readiness-levelmanufact
http://aviation-africa.eu/sites/default/files/events/105%20SIASA%20-
%20Introduction%20Fly%20By%20Wire%20Aircraft%20and%20New%20Technology.pdf
http://rockwellcollins.com/Our_Company.aspx
http://rockwellcollins.com/Our_Company.aspx
http://saab.com/commercial-aeronautics/aircraft-support-solutions/
http://www.airbus.com/uploads/pics/Material_andLogistics_intro_section__1_.png
http://www.businessdictionary.com/definition/program-management.html
http://www.davi.ws/avionics/TheAvionicsHandbook_Cap_11.pdf
http://www.dodmrl.com/
http://www.evmlibrary.org/library/TP-16%20Schedule%20Analysis%20Techniques,%20Majerowicz.pdf
http://www.garmin.com/en-US/company/about/
http://www.garmin.com/en-US/company/about/
http://www.herouxdevtek.com/company/corporate-profile
http://www.herouxdevtek.com/company/corporate-profile
http://www.kollmorgen.com/en-us/service-and-support/partners/supplier-terms-conditions/
http://www.moog.com/products/actuators-servoactuators/aircraft/flight-control/
http://www.pcubed.com/services/glossary.program
http://www.pmknowledgecenter.com/dynamic_scheduling/risk/schedule-risk-analysis-how-measure-
your-baseline-schedule%E2%80%99s-sensitivity
http://www.price-induction.com/about-us/who-we-are/
http://www.price-induction.com/about-us/who-we-are/
https://buy.garmin.com/en-US/US/in-the-air/flight-decks/g1000-/prod6420.html
https://buy.garmin.com/en-US/US/in-the-air/flight-decks/g1000-/prod6420.html
https://www.law.cornell.edu/cfr/text/48/7.105
https://www.nasa.gov/content/technology-readiness-level/#.VTnRPSFViko
https://www.rockwellcollins.com/Products_and_Systems/Controls.aspx
https://www.rockwellcollins.com/Products_and_Systems/Controls/Fly_by_wire.aspx
https://www.rockwellcollins.com/Products_and_Systems/Controls.aspx
https://www.rockwellcollins.com/Products_and_Systems/Controls/Fly_by_wire.aspx
Rockwell Collins. (2015). PFCC-4100 Primary Flight Control Computers.
http://www.rockwellcollins.com/Data/Products/Controls/Fly-by-
wire/Primary_Flight_Control_Computers.aspx