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Conceptual Design Review Report

Team 3 Tom Zettel

Mike Bociaga Jonathan Olsten

Jamie Rosin Hayne Kim

Brandon Washington Marques Fulford

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Table of Contents Mission Statement 3 Mission Plans 3 Walk Around 3 Advanced Technologies 4 Carpet Plots 8 Sizing Approach 9 Design Trade-Offs 12 Interior Layout 15 Aerodynamics 17 Propulsion 23 Performance 25 Structures 29 Weights and Balance 30 Terminal Servicing 36 Enviromental Impact 38 Maintenance and Reliability 39 Summary 40

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Mission Statement

To create an innovative and cost effective commercial aircraft capable of take-off and landing in extremely short distances, making it available to a larger number of runways, in order to open up more airports, primarily to relieve the continuous growing congestion of large hubs.

Mission Plans The goal of Team Arrival’s aircraft is to relieve congestion at major hubs. The

three hubs that experience the most congestion (according to www.bts.gov) are Chicago O’Hare, New York LaGuardia, and Newark International. The first mission is to take off from Gary Chicago and land at Dallas Love Field. Gary Chicago airport is located 42 miles from Chicago O’Hare and Dallas Love Field is located 20 miles from Dallas International. This mission is intended to redirect some of the traffic from both Chicago O’Hare and Dallas International to secondary airports that experience less congestion. Team Arrival’s aircraft will have the capability of takeoff and landing on runways of 3000 feet or less making this mission possible.

The second mission is to do a half-runway takeoff from New York LaGuardia and do a non-interfering spiral descent at Miami International. Both of the airports at LaGuardia and Miami have a significant amount of road traffic around them and it is Team Arrival’s feeling that most passengers will not want to fly into a secondary airport and then sit in street traffic for an extended period of time in order to get to their final destination.

The third mission is to do a half-runway takeoff from Charlotte International, land at Essex County, then takeoff from Essex County without refueling, and return to Charlotte International and do a non-interfering spiral decent. This mission was planned as a round trip without refueling because Essex County is currently a small airport that may not have the equipment or personnel to refuel the aircraft and get it ready to takeoff again in a reasonably short amount of time. If this airport sees an increase in use in the future and increases it’s equipment and personnel this mission can be changed to refuel at Essex County. It would be more cost effective to refuel the aircraft because the aircraft would be lighter upon initial takeoff. This is hopefully something that can be done in the future, but Team Arrival is preparing for the current conditions.

Walk-Around A walk around view of the aircraft can be seen in Figure (1) below. The walk around showcases the advanced technologies that were integrated into the design concept. The forward swept wing was added in order to aid in the aircraft’s ESTOL capability. The composite fuselage was meant to decrease the overall weight of the aircraft. The upper surface blowing was designed to increase the CLmax of the aircraft. The geared turbofan was integrated in order to make the aircraft more fuel efficient. The advanced integrated flight deck was added to increase safety of the aircraft. The extra large cabin

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windows were thrown in to increase passenger comfort as an added bonus. For full dimensions and other views of the aircraft please refer to Appendix (A).

Figure (1) Walk around of aircraft concept

Advanced Technologies Forward Swept Wings (FSW)

One of the main focuses of the design of the aircraft was minimizing the takeoff distance. FSW were chosen primarily because of their ability to assist in a shortened takeoff distance. The FSW produces more lift at takeoff speeds than does a traditional wing. Another important feature for reducing takeoff distance is that the FSW stalls at a higher angle of attack and therefore allows for a greater rotation angle on takeoff. Secondarily, the forward swept wing allows for an increased wing thickness-to-chord ratio; this will increase fuel volume and result in a wing weight reduction which would in turn lower acquisition costs.

While the FSW offers the above advantages, there is a salient problem when using FSW; aeroelastic divergence occurs at lower mach numbers than with an aft swept wing. A reduction in divergence speed on the order of 90% can be expected when a wing is swept forward from 0 to 28 degrees. Although it will remain a fact that the divergence dynamic pressure of swept-forward designs will always be lower than their unswept counterparts, the important consideration that designers face is associated with the amount of additional structural weight needed for the increased stiffness required to insure the absence of divergence within the operating performance envelope of the aircraft . This weight penalty is known to be very severe for conventional metal wings.

With the unique properties of advanced composites, the basic material properties can be tailored to suit a particular loading condition. This additional design parameter is the key to controlling and effectively adjusting the wing stiffness characteristics to combat the weight problem caused by the lower divergence speeds associated with swept-forward wings.

Geared

Turbofan Engine

Forward Swept Wing

Upper Surface Blowing

Composite Fuselage

Advanced

Integrated Flight Deck

Extra Large

Cabin Windows

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Composites Weight Savings

The primary advantage of increasing the amount of composite materials on the aircraft is to decrease its overall weight. Reduced weight ultimately leads to reduced operating costs and can save airlines billions of dollars over the life of the aircraft. The Boeing 787's all-composite fuselage makes it the first composite airliner in production. The new airplane will be composed of 15% aluminum, 50% composite (mostly carbon fiber reinforced plastic) and 12% titanium. Each fuselage barrel will be manufactured in one piece, and the barrel sections joined end to end to form the fuselage. This will eliminate the need for about 50,000 fasteners used in conventional airplane building. The superior strength of the composite fuselage will allow higher pressurization in the passenger cabin, making it easier to control temperature, humidity and ventilation.

Figure (2) Capabilities of composite material

By comparing the empty weight fraction of Boeing’s 787 to similar size aircrafts

such as Boeing 777 and Airbus A330, it is possible to find the trend weight savings as shown in Figures (3) and (4). Following this trend it is estimated that the aircraft will have a 15% weight savings factor on the empty weight compared to current similar size aircrafts. This weight saving can also be the solution to forward-swept wing’s divergence phenomenon.

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Figure (3) Empty Weight Fraction Comparison

Empty Weight Fraction Material Comparison

y = 233.08x-0.5017

y = 851.58x-0.6321

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0 50000 100000

150000

200000

250000

TOGW [lb]

Em

pty

Wei

gh

t F

ract

ion Boeing/Airbus Current

Boeing/Airbus CFRP

Power (Boeing/AirbusCurrent)

Power (Boeing/AirbusCFRP)

Figure (4) Empty Weight Fraction Material Comparison

Following the trends shown above for an aircraft with 50% composites by volume, it is thought that in 2058 an aircraft with 80% composites by weight will be feasible based on the work Boeing is currently doing. Therefore, the overall weight savings sought by using the composite materials incorporated into calculations is 25%.

Empty Weight Fraction Comparison

y = 3853.1x-0.6833 y = 549.16x-0.5221

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0 200000 400000 600000 800000 1000000

TOGW [lb]

Em

pty

Wei

ght F

ract

ion

Airbus A330

Boeing 777

Boeing 787

Power (Boeing 787)

Power (Boeing 777)

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Upper Surface Blowing Upper Surface Blowing (USB) is another technology that is primarily in place to

increase flight performance and decrease takeoff distance. As can be seen in Figure (5) below, the primary difference in USB is that the engines are mounted above the wings as opposed to the traditional mounting of engines below the wings.

Figure (5) Depiction of Upper Surface Blowing

In the 1970’s this technology was put into place on the first models both the Boeing YC-14 and the Antonov An-72 because of its ability to create additional lift and thus decrease the takeoff distance of these aircraft. Since that time, upper surface blowing has been in limited use on several different aircraft because of its ability to significantly decrease takeoff and landing distances and to increase cruise and max CLmax.

Team Arrival’s aircraft utilizes USB and its affects are taken into consideration in the calculation of CLmax and also the approximated takeoff and landing distances

Specific Fuel Consumption

With the consistent increase in fuel prices and the expectation that this trend will continue, one vital portion of success in the aircraft industry is specific fuel consumption. The following research was done by Bombardier on SFC trends amongst un-ducted fan and geared turbofan engines.

Power Series Projection

Certification

Year SFC

UDF SFC

(15%

savings)

UDF SFC

(25%

Savings)

GTF SFC

(15%

Savings)

GTF SFC

(20%

Savings) 2010 0.59 0.51 0.45 0.51 0.48 2015 0.57 0.49 0.43 0.49 0.46 2020 0.56 0.47 0.42 0.47 0.44 2025 0.54 0.46 0.40 0.46 0.43 2030 0.52 0.44 0.39 0.44 0.42 2035 0.50 0.43 0.38 0.43 0.40 2040 0.49 0.41 0.36 0.41 0.39 2045 0.47 0.40 0.35 0.40 0.38 2050 0.46 0.39 0.34 0.39 0.36

Table (1) Power series projection In addition to the research done by Bombardier, in order to develop an accurate

projection of the aicraft’s SFC, Team arrival has also compiled data on SFC of current

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aircraft fleets as well as SFC data from aircraft that are exiting service to determine the trend in SFC over the next 50 years. The results of this process can be seen below:

Figure (6) Plot of SFC versus certification date

Based on the precision between both of these SFC projections, Team Arrival

expects that a geared turbofan in service in the year 2058 will have an SFC of 0.36. This will not only significantly decrease operating costs, but will also increase possible mission range for the aircraft.

Carpet Plots The final carpet plot used to pick the design point of the aircraft can be seen in Figure (7) below. The carpet plots were designed by inputting the wing loading and thrust-to-weight values into the sizing code and receiving the gross takeoff weight as an output. These values were then plotted resulting in the tilted square shape. The takeoff constraint line was added by first plotting the takeoff distances against wing loading for various thrust-to-weight values. The point at which the wing loading matched the takeoff distance constraint of 1500 feet was recorded and added to the overall carpet plot. This resulted in the takeoff distance constraint which ended up being the most restricting of all the constraint values. The rest of the constraint lines did not play a factor in deciding the design point. The next most restricting constraint was the velocity at best range, which can be seen in Figure (7), and clearly it does not effect the design point. The point at which all of the constraints are met and a maximum wing loading can be achieved is with a wing loading of 84 and a thrust-to-weight value of 0.310, which is the overall design point for the aircraft.

SFC vs. Certification Date

y = 2E+44x-13.476

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

1960 1980 2000 2020 2040 2060 2080

Certification Date

SF

C

SFC

Power (SFC)

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Figure (7) Carpet plots used to choose design point for aircraft

Sizing Approach The sizing approach used by Arrival for the aircraft used methods from Daniel P. Raymer’s book, “Aircraft Design: A Conceptual Approach”. This approaches included initial sizing based off of empty weight fraction and then using component weight breakdown for detailed weight estimation. To start off, Arrival used some fixed design parameters for the empty weight fraction code “sizing.m” which can be seen in the “Weight and Balance” section. The fixed design parameters used for sizing of the aircraft are given in Table (2).

FFiixxeedd PPaarraammeetteerrss VVaalluuee CCLL,,mmaaxx

44 tt//cc 00..1111

ΛΛ aatt ..2255cc ((ffoorrwwaarrdd)) 2255°° λλ 00..227788

AARR 1100 Table (2) Fixed Design Parameters for Arrival’s Aircraft

These parameters came about as a result of the basic constraints and the mission scenarios that the aircraft is designed around such as Extremely Short Take-Off and Landing and a target range mission of 2000 nmi. Each mission scenario was broken up into weight fractions based off of mission phase, such as takeoff, climb, cruise, and

T/W = 0.283

T/W = 0.303

T/W= 0.323

W/S = 64 [lb/ft2]

W/S = 74[lb/ft2]

W/S = 84 [lb/ft2]

Velocity Best Range

90600

90700

90800

90900

91000

91100

91200

91300

91400

91500

91600

W0

[lb

]

W/S = 84 lb/ft2

T/W=0.310

Ground Roll <= 1500ft

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landing. Each mission also used its specified range which would either be maximum target, maximum threshold, or the distance between city pairs. Other values, such as cruise Mach, cruise L/D, 177 passengers and 6 crew (2 cockpit, 4 flight attendants), and initial thrust-to-weight (T/W) and wing loading (W0/S) estimates are also used as inputs. The code sizes the aircraft and calculates design weights based off of range and payload requirements. It iterates by taking the design weight value and subbing the guess weight value until the design takeoff weight agrees with the guess weight value. These values are then used to create carpet plots which allowed us to select the aircraft final design point T/W and W0/S. With these values the code is run again giving the final weight estimates and initial dimensional sizing of the aircraft’s wing, engine, and tail surfaces. These weight estimates and initial dimensions were then used as inputs in the weight breakdown code, “weight_breakdown.m”, which gives the final weight values for the whole aircraft, component weights, and static margin results. If the weight values from the weight breakdown code were in disagreement with the initial “sizing.m” code, then the sizing.m code would be rerun with new factors in the empty weight equation to try to match the weights of the weight breakdown code, which Arrival assumes is more accurate since it is based off of specific component weights rather than the whole aircraft. The process is then iterated until all the weight values reasonably agree, resulting in final weights, dimensions, and static margin values. A flowchart for this process is provided in Figure (8).

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Figure (8) Sizing Approach Flowchart

For sizing of the tail surfaces, Arrival considered various scenarios as well as some fixed tail parameters. The fixed parameters for the Horizontal Tail can be seen in Table (3), while the fixed parameters for the Vertical Tail can be seen in Table (4).

HHoorriizzoonnttaall TTaaiill PPaarraammeetteerrss VVaalluuee

MMoommeenntt AArrmm [[ ff tt]] 4400

AARR 55..5544

λλ 00..118866

ΛΛ aatt ..2255cc 3300°° Table (3) Fixed Design Parameters for aircraft horizontal tail

YYeess

NNoo

Input: Fixed Design

Parameters

Size Aircraft

Weight, Surfaces, Fuel

Empty Weight

based on Component

Weight

Empty

Weights

Final Sizing Results

Empty Weight

Fraction Estimation

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VVeerrttiiccaall TTaaiill PPaarraammeetteerrss VVaalluuee

MMoommeenntt AArrmm [[ ff tt]] 4400

AARR 11..5566

λλ 00..3311

ΛΛ aatt ..2255cc 3355°° Table (4) Fixed Design Parameters for aircraft vertical tail

These parameters along with scenarios such as single engine failure during takeoff, crosswind, and 2.5g maneuver at altitude were used to size the vertical and horizontal tails. The horizontal tail also needed to be able to provide the proper moment to ensure a high enough takeoff rotation angle. Another consideration for tail sizing was the moment arm length from the quarter-chord of the main wing. The closer to the main wing the tail surfaces were, the larger they would become in area. The affect was also felt with wing area. The lower the wing loading, the higher the wing area, particularly with our high aspect ratio wing, and thus, the larger the tail surfaces need to be to counter the moments from the wing. As a result, final configuration was important for final sizing. The final dimensional sizing results are given in Table (5).

Aircraft Dimensions Value

Wingspan [ft] 102

Wing Area [sq. ft] 1,032

Wing Mean Aero. Chord [ft] 11

Vertical Tail Height [ft] 36

Vertical Tail Area [sq. ft] 560

Horizontal Tail Span [ft] 66

Horizontal Tail Area [sq. ft] 450

Fuselage Length [ft] 123 Table (5) Dimension Sizing results for aircraft

Design Trade-Offs The aircraft’s configuration changed from SDR as the design was adjusted to address difficulties encountered during the design process. The original design of the aircraft called for a forward swept wing mounted as aft as possible – near the aircraft’s vertical tail - with a lifting canard mounted as far forward as possible on the nose. Both of these surfaces were mounted below the fuselage. In addition, two geared turbo-fan (GTF) engines were mounted on the upper surface of the wings to allow the aircraft to

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take advantage of upper-surface blowing (USB) technology which redirects engine exhaust flow over the surface of the wing and flap using the Coanda effect, thus increasing the maximum achievable lift coefficient. The advantages of this configuration were that with a forward swept wing, the wingbox could be located near the tail. With the box near the tail, the landing gear would be as far aft as possible while not having to be very tall thanks to the high mounted USB engines. This gave two immediate advantages, one being that the gear further aft allowed for a higher angle of attack on takeoff rotation, and the other being that the height from the ground to the door sill could be kept to a minimum. The higher angle of attack on rotation is useful since a forward swept wing better maintains stability and has a lower stall speed even at high values of α. The canard at the nose of the airplane had a long moment arm to assist in rotating the aircraft to its high rotation angle. The rotation angle for this configuration was 18°. Figure (9) shows the original SDR configuration of the Arrival aircraft.

Figure (9) Arrival aircraft configuration for SDR

As the design process continued, it became apparent that this configuration had a few challenges that needed to be addressed. The first issue was with the low-mounted canard and “ramp rash”, a term describing the tendency of ground crew to unintentionally damage the aircraft with ground equipment. Another issue was that airflow into the engines would be disturbed by the lifting canard if the current configuration was kept. As a result the first change was to move the canard to the top of the aircraft. However, with the curvature of the nose of the fuselage, the canard would have to be moved further aft to eliminate interference with cockpit space, the passenger cabin, or the aisle ceiling. Furthermore, jet-bridge compatibility would require the main loading door be in front of the canard to avoid interference with the jet-bridge. As a result of the shortened canard moment arm, the size of the canard began to grow. Similarly, with such a short moment arm between the CG and vertical tail, aircraft controllability required an enormous

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vertical tail – with sizes on the order of the 747s vertical tail size for our 737-sized aircraft. Figure (10) shows the current configuration of the Arrival aircraft.

Figure (10) Arrival aircraft current configuration

One consideration for the current configuration of the aircraft was keeping the horizontal tail surface from being in the wash of the upper-surface mounted engines. However, a quick check of the projected exhaust path of the engines versus the location of the horizontal tail showed that as long as the horizontal surface is mounted as high up on the fuselage as possible, this wouldn’t be an issue. Not wanting to incur the weight penalties from a T-tail design, the horizontal tail is mounted on the vertical tail in a cruciform fashion, but is so low on the vertical tail, that the whole system can be treated as a conventional tail. While the horizontal tail will have an elevator, control analysis showed that a variable incidence tail would be desirable for trim. With the horizontal stabilizer redesigned, the wing was moved forward to about 75% of the length of the fuselage as measured from the nose to the wing quarter-chord. This allowed the moment arm for both the vertical tail and the horizontal tail to be lengthened, thus allowing those surfaces to be reduced in area. Moving the wing further forward was considered, however, in order to maintain a high takeoff rotation angle, the landing gear would have to be lengthened. A forward swept wing still allows the wingbox to be located further aft on the aircraft. Although an additional 4° of takeoff rotation angle was lost, the trade-off increased design simplicity, facilitated easier placement of the emergency exits, as well as simplicity of manufacturer’s learning-curve development with a more conventional layout. However, with lightweight CFRP materials, efficient GTF engines, a forward swept wing, and lightweight fly-by-light fiber optic data cables, the aircraft still has a major technological advantage over contemporary single-aisle aircraft. All design trade-offs that Arrival made were based around the target requirements for the aircraft as well as general functionality. Most important among them was the

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ESTOL requirement. For example, a high aspect ratio wing with low wing loading was desirable for the purpose of meeting the 1500 ft ground roll requirement of the aircraft. However, a wing loading below 80 psf would result in the aircraft being unacceptably vulnerable to turbulence at altitude. As a result, Arrival expanded its carpet plot analysis and increased the thrust-to-weight ratio of the aircraft to allow for a higher wing loading of 84 psf, thus reducing the effect of turbulence on the aircraft. For comparison, a Boeing 737-800 has a wing loading of about 115 psf, however the 737 also has a much longer takeoff distance.

Interior Layout The interior of the aircraft was modified after the SDR presentation in order to be more practical than the previous layout. The original economy only layout shown below in Figure (11) called for 4 lavatories, 2 galleys (one of which was quite large), and 1 closet. The two class layout, shown below in Figure (12) had the same arrangement of facilities, but the first 5 rows are replaced by 4 rows of first class. The final economy only layout is shown below in Figure (13), and the final two class layout in Figure (14).

Figure (11) Preliminary Economy Only Layout

Figure (12) Preliminary Two Class Layout

Figure (13) Final Economy Only Layout

Forward

Rear

Rear

Forward Close

Emergency Exit Rows

Close

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Figure (14) Final Two Class Layout

For the final layout the amount of galley space was reduced from the preliminary designs. This decision was made due to the current trend of reduced service on airlines to help offset rising costs of operation. Both the economy and two class layouts feature a fore and aft galley of equal size which will provide drink, snack, and limited hot meal service to the first class passengers. The number of lavatories was also reduced, as the final layout has three compared to the original four. The size of the lavatories was also reduced, because the preliminary lavatories were much larger than what is standard on a narrowbody aircraft. Both layouts have 2 rear lavatories, and one forward lavatory which would be exclusive to first class passengers in the two class layout. The final item on the floor plan is a small closet behind the forward galley. The seating layout has seen only minor changes since the SDR presentation. The economy seats are 18 inches in width, with 32 inches of seat pitch which are average values for economy seating in current aircraft. The first class seats are 20 inches in width with 38 inches of seat pitch. All of the seat dimensions were compared to statistics taken from seatguru.com. It should be noted that there are large variations in the values for the first class seat pitch than the economy. Removing the extremely high values (50+ inches) gives an average seat pitch of 37.5-38 inches, which agrees with what was chose for the aircraft’s seating. A slight increase in seating capacity was made due to the addition of an extra row of economy seats on the starboard side in both configurations. In the economy only layout there are 29 full rows with six seats each and then the additional half row with three seats for a total seating capacity of 177 passengers. The two class layout has 4 rows of seats with 4 seats in each row for a total capacity of 16 first class passengers. The economy section has 24 full rows of seats, and the half row with the same configuration as the economy only layout. This gives an economy capacity of 147 passengers, which yields a total aircraft capacity of 163 passengers. The main exit locations haven’t changed since the preliminary designs. The final design has a pair of doors in the front and rear of the plane denoted by the yellow arrows on Figures (13) and (14). Terminal servicing calls for passengers to be loaded and unloaded through the port front door. To meet safety requirements for a timely evacuation in an emergency there are 2 emergency exit rows located by the red sections in Figures (13) and (14). These rows will have an increased seat pitch of 36 inches to allow for an increased flow of people through the exits in the event of an emergency.

First Class

Economy

Economy

First Class Galley

Close

Emergency Exit Rows

Closet

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Both rows will make use of emergency slides for evacuation, and the over-wing exit will require a special heat resistant slide due to the upper surface blown engines.

Aerodynamics The most important aspect of the aerodynamic design of the aircraft is airfoil selection. The wing needed for an aircraft that cruises at Mach 0.78, at an altitude of about 36,000 ft Mean Seal Level (MSL), must also be optimized for a 2000 nmi maximum range, and be able to meet the ESTOL requirements Arrival is designing the aircraft to meet. The wing design will likely occur in the 2040s for a service entry of 2050 to 2058. In the meantime a representative airfoil was chosen based off of Arrival’s needs to show what this future wing will need to be capable of doing. The airfoil is a supercritical airfoil from the Gulfstream GIII aircraft found on the University of Illinois Urbana-Champaign (UIUC) airfoil database created by Michael Selig.

A supercritical airfoil was chosen to allow the aircraft to cruise at a higher Mach number without incurring wave drag or early separation due to shocks forming at critical Mach number. In other words, the point is to raise the critical Mach number of the aircraft. The GIII airfoil sections were selected since the GIII cruises at about the same speed for the same range as the Arrival aircraft. Figure (15) shows a GIII airfoil section from about the midspan of the GIII wing. Figure (16) shows the lift curve of a mid-span airfoil section for the GIII when analyzed using a Potential Flow code. It is important to note that because the analysis used was potential flow, a reliable Clmax cannot be determined for the airfoil. A viscous analysis must be performed instead.

Figure (15) GIII BL167 Airfoil Section

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Figure (16) GIII BL167 Airfoil Section Lift Curve Slope

For the tail section airfoils, Arrival looked at the performance need of the horizontal and vertical stabilizer to choose the best representative airfoil. The vertical tail will have a symmetric airfoil, since the only time it needs to provide a side force is when the rudder is defected creating a camber. For the horizontal tail, all that is necessary is for the surface to counter the down pitching moment of the main wing with the elevator deflection at 0°. As a result, a symmetric airfoil has been selected for the horizontal tail as well. Both tail surfaces use an airfoil section with about 12% thickness. Aerodynamic changes to the tail may happen as Arrival progresses with the design towards the date of entry into service of the aircraft. Another consideration for the aircraft’s wing design is takeoff performance, which will largely be impacted by CLmax. High-lift devices are a major part of the CLmax equation. For the section of the wing that has USB, the flap segment must allow the Coanda effect to take place. As a result, there can be no slots, but rather the flap has to provide for an uninterrupted upper surface. As a result, plain flaps were selected for the inboard section of the aircraft’s wing in the area behind the engines. Figure (17) shows the flap set-up for a USB engine.

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Figure (17) Flap Set-up for a USB engine, in this case, the Boeing YC-14

For the rest of the flapped area of the wing, double-slotted Fowler flaps were selected where the USB effect is negligible. The Fowlers were chosen for there high ∆CLmax contribution to the overall CLmax of the aircraft. Figure (18) shows the typical set-up for a double-slotted Fowler flap system.

Figure (18) Flap Set-up for a double-slotted Fowler flap system

The flaps were sized based off of the aircraft’s wing mean aerodynamic chord and span. A certain portion of the span was allocated for the ailerons which will be sized as Arrival progresses with the design. The preliminary flap sizing along with the preliminary CLmax calculation are found using the “High_Lift.m” code found in Appendix (B) which is based off of flap-sizing principles found in Daniel Raymer’s text. USB is expected to provide to the CLmax in a huge way according to research found from the AIAA paper titled “Advanced Circulation Control Wing System for Navy STOL Aircraft.” (AIAA-57598-949). From that paper, we get Figure (19) which plots the CLmax of a twin-engined USB aircraft against its aspect ratio.

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Figure (19) CLmax vs. AR for a twin engined STOL aircraft per AIAA-57598-949

Using this data, Arrival can assume that USB will provide a large portion of the CLmax necessary to meet the ESTOL takeoff requirement. Arrival’s target CLmax is 4. The wing alone, with plain flaps and double slotted Fowler flaps provide a CLmax of 1.85. Arrival will continue to refine the wing design too provide a higher CLmax value. In the meantime, USB can be assumed able to make up for the shortfall of 2.15. Since the current wing is only a representative one from Gulfstream aircraft, Arrival’s future wing design will have a better CLmax value. This will ensure that landing performance is not adversely affected by the USB engines being throttled at idle during approach and landing. Drag Buildup

In order to best estimate the aircraft drag, the parasite drag was estimated using a “drag buildup” approach. In such an approach, the drag of individual components were approximated and combined to account for skin friction, drag interference and separation drag using skin friction coefficients, interference factors and form factors, respectively. The total parasite drag for the aircraft was computed using Equation (1).

misc

c

Dref

wetccfc

D CS

SQFFCC +=

∑0

(1)

Five major components were taken into consideration for the clean configuration:

the fuselage, wings, horizontal tail, vertical tail, and engine nacelles. In order to compute

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skin friction coefficients, Equations (2) and (3) were used to determine the skin friction on a flat plate at the same Reynolds number as each of the five major components. The Reynolds numbers used were the lower of the two between the Reynolds number and the cutoff Reynolds number computed in Equation (4). For all five components, the cutoff Reynolds number was significantly larger than the Reynolds number; this was because the smoothness (k) is very low for smooth, molded composites.

Re

328.1=fC (Laminar) (2)

( ) ( ) 65.0258.2 144.01Relog

455.0

MC f

+= (Turbulent) (3)

( ) 16.1053.1Re Mk

LCcutoff = (4)

Where C = 38.21 for subsonic flight, and C = 44.62 for transonic flight To determine the effective skin friction over each component, laminar flow was

assumed over certain percentages of each surface. These percentages are shown in Table (6) and represent a significant improvement of the state of the art in terms of how well laminar flow is maintained across surfaces. The Piaggio GP180 was able to achieve 50% laminar flow over the wings and tails and 20-35% over fuselagei. Better aerodynamic tailoring for laminar flow, improvements in composites manufacturing and smoothness will surely increase the amount of laminar flow maintained across surfaces of aerospace vehicles in the future.

The Form Factors, Interference Factors, and Skin Friction Coefficients computed or assumed for each of the five components are shown in Table (6) for the clean configuration at 36,000ft altitude. A moderate reduction was applied to the Form Factors to account for future improved aerodynamic efficiency. When applied as in Equation (1), these coefficients and approximations for the wetted areas of each of the components produce an estimated parasite drag value of approximately 0.0136.

Aircraft Component

Form Factor

Interference Factor

Skin Friction Coefficient

Percentage Laminar

Fuselage 1.044 1.00 0.001036 60% Wing 1.215 1.00 0.000294 70%

H. Tail 1.229 1.04 0.000269 70% V. Tail 1.240 1.04 0.000355 70% Nacelle 1.082 1.25 0.000255 50%

Table (6) Parasite Drag Buildup Drag Polar

Three components of drag contribute to the total aircraft drag: parasite, induced, and wave drag. In order to most usefully represent the drag coefficient of the aircraft, one may generate a drag polar. The parasite drag buildup was discussed in the previous section. Induced drag was determined using the theoretically-based Equation (5) below. Wave drag was modeled using an empirical piecewise continuous curve-fit, which can be

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found in “drag.m” in Appendix (B). Figure (20) below shows wave drag as a function of Mach number; this figure shows that the model for wave drag used follows the correct trend for wave drag even though it is not a direct theoretical solution.

eAR

CC L

iD π

2

= (5)

Figure (20) Wave Drag as a Function of Mach Number

The drag polar combines all three categories of drag as a function of lift

coefficient. Equation (6) describes the coefficient of total drag as the sum of its three components. The drag polar is shown in Figure (21) for the aircraft in the clean configuration at a cruise Mach of 0.78.

DwiDDD CCCC ++= 0 (6)

23

Figure (21) Drag Polar in Clean Configuration (M = 0.78)

Propulsion Description

A geared turbofan was selected as the rubber engine type of choice for Team Arrival. A geared turbofan uses a gearing system to increase the efficiency of the fan by running at an optimal fan RPM at different turbine RPMs. This allows for a higher bypass ratio as the fans diameter can increase while still maintaining subsonic fan tip speeds. Its increased bypass ratio over conventional turbofans helps increase fuel economy and extends range; the justification of the improved specific fuel consumption is explained in “Advanced Technologies” section. The increased fuel economy greatly reduces the MTOW by allowing for much lower fuel weight fractions, requiring less aircraft to carry the reduced fuel weight. A cutaway view of a geared turbofan engine is shown in Figure X.

Figure (22) Geared Turbofan Cutaway Viewii

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The engine will be mounted above the wing to make use of the upper surface blowing discussed in “Advanced Technologies” sections. The wing-mounted configuration of the engine provides bending moment relief to the wing spars, but does force the vertical tail to be larger than if they were mounted on the fuselage; this is a result of needing to provide enough yaw control to counter the moment caused by one engine being out and is discussed in greater detail in “Tail and Control Surface Sizing” section.

This engine configuration also affects maintenance and reliability. Since it is mounted forward of the wing to provide maximum upper surface blowing effect, it is still easily accessed from the ground; the engine is mounted higher but can still be lowered straight out of the nacelle. Furthermore, with fewer compressor stages, maintenance of the geared turbofan is also cheaper; Pratt & Whitney cites maintenance costs on the P&W 8000 as 30% less than similar-thrust regional and single-aisle jet engines because of this fact. P&W also believes that current geared turbofan technology reduces engine noise levels by as much as 30 dBiii .

The Pratt & Whitney 8000 geared turbofan was used as a baseline for the analysis of a futuristic “rubber” geared turbofan. The PW8000 was designed as a more economic engine for regional and single-aisle passenger jets. The specifications for the both the P&W8000 and future geared turbofans, as well as the next generation 737 engine (CFM56-7), are seen in Table (7). The future geared turbofan specifications are estimated by the thrust required for a single-engine takeoff, future weight savings and improvements in efficiency.

CFM56-7 P&W 8000 Future GTF Thrust (lbs) 26400 30000 28900 Weight (lbs) 5256 8000 5000 Bypass Ratio 5.1 11 12 Low P Comp. Stages 5 4 4 High P Comp. Stages 9 5 5 Diameter (in) 61 76 65 Length (in) 99 124 99

Table (7) Engine Specifications

In order to determine available thrust and drag at important points in the aircraft’s mission, the drag was estimated as discussed in the previous section, and the engine thrust was estimated using the approach found in Hill & Petersoniv. In this analysis, the decrease in thrust with altitude was a function of three things: ambient air stagnation pressure, air density, and ram drag. Equation (7) is the generic equation for turbofan thrust and was used in this approach. Any bleed air that might be drawn off the engines was neglected in this analysis.

( ) ( ) ∞+−++= UUUfT efe ββ 11 (7)

where f is the fuel-to-air ratio, β is the bypass ratio, and U is velocity at the core exit, fan exit, and free-stream.

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The thrust-drag-velocity diagram was determined at two important cruise altitudes for the clean configuration. Figures (23) and (24) depict this diagram for 30,000 and 36,000 ft altitude, respectively. One can see that the engine thrust lapses with altitude and airspeed, as is expected with gas-turbine engines. Drag also has the expected curvature for flight across this velocity range.

Figure (23) T-D-V Diagram (30,000 ft) Figure (24) T-D-V Diagram (36,000 ft)

Performance Having defined the dimensions and weights of the aircraft, it was possible to

determine the safe flight limits of the aircraft through the flight envelope and the V-n Diagram as discussed in the following subsections. Flight Envelope

The aircraft flight envelope maps the combinations of altitude and velocity that the aircraft has been designed to withstand. The excess power available to the aircraft is the main factor in determining both the absolute ceiling as well as the service ceiling. Excess power, Ps, is defined as shown in Equation (8).

22

02

1( ).5 R e

/ .5s l D

s l

nT V C WAP s VW W S V S

ρρ πρ ρ

= − −

(8)

ρ=Air density T = Thrust W/S = Wing Loading AR = Aspect Ratio V = Velocity CD0 = Zero Lift Drag coefficient

Using the above equation, the absolute ceiling was determined by varying the

velocity with zero excess power. Similarly, according to FAR Part 25, the service ceiling

Vbest range M = 0.78

26

was found by varying the velocity and setting the excess power to 500 ft/min. In order to calculate the stall limit the VStall Equation (9) below was used.

2Stall

MaxL

WV

SCρ= (9)

In this instance, the altitude was input indirectly through variation of the density

and then the corresponding velocity was found. Using these three boundaries, the flight envelope was used to determine the safe limits for operation of the aircraft. This flight envelope can be viewed in Figure (25) below.

Figure (25) Plot of flight envelope

Once the boundaries for the flight envelope were found, the cruise altitude and cruise speed were chosen within the given stall and service ceiling parameters and can be seen above.

V-n Diagram

Strength requirements for an aircraft are specified in terms of limit loads and ultimate loads. Limit loads are the maximum loads expected to be experienced by the aircraft during service and ultimate loads are the limit load multiplied by a factor of safety. The factor of safety is determined by the FAA and it is required that all aircraft are capable of meeting these criteria for airworthiness certification. Therefore, the Federal Aviation Regulations become the driving factor in determining structural stability

Flight Envelope

0

5000

10000

15000

20000

25000

30000

35000

40000

45000

50000

55000

0 100 200 300 400 500 600

Speed (knots)

Alt

itu

de

(Ft)

Absolute Ceiling Stall Limit Service Ceiling

Cruise Altitude Cruise Speed : 447 knots

27

at given speeds. The FAR states that the limit load factor, n, be defined by Equation (10) below.

24,0002.1

( 10,000 )

lbn

W lb= +

+ (10)

W=Design Maximum Takeoff Weight

With a design maximum takeoff weight of 86,800 lbs, the load factor n was 2.35,

however, Part 25 of FAR states that if the limit load factor is not to be less than 2.5. Therefore, in order to ensure that the aircraft would be able to maneuver under a load factor of 2.5, several similar sized single aisle aircraft were examined and were found to be within compliance with the FAA regulations. It is therefore assumed that using the advanced materials and control surface operations available in 50 years, the aircraft will be able to meet these same FAA regulations. The V-n diagram showing the loads encountered at various flight speeds is seen in Figure (26) below.

Figure (26) Plot of V-n diagram

Performance Velocities

In order to calculate the best range velocity of the aircraft, equation 17.25 from Raymer was used, this can be seen in Equation (11) below.

28

0

132 Re

BestRangeD

W AVS C

πρ

= (11)

ρ=Air density W/S = Wing Loading AR=Aspect Ratio e=Oswald Efficiency Factor CD0 = Zero Lift Drag coefficient

The best endurance velocity is the velocity at which the aircraft stays in the air the longest. It is also the velocity at which the power required and thus the fuel flow to the engine is at a minimum. Typically, the best endurance velocity was determined to be approximately 75% of the velocity for best range. The maximum speed was chosen by examining the thrust and drag at various airspeeds as seen in Figure (27) below.

Figure (27) Plot of thrust and drag at different airspeeds

Best Range Velocity 482 knots Best Endurance Velocity 361 knots Maximum Speed 488 knots

Table (8) Aircraft performance characteristics

The takeoff stall speed was calculated using the stall equation seen in Equation (12) below.

29

2

Stall

MaxL

WV

SCρ= (12)

The density at 5,000 feet was used in order to represent a worst case takeoff scenario. Once the stall speed was calculated, the takeoff speed is then taken to be 110% of the stall speed and the approach speed is 115% of the stall speed.

The takeoff and landing distances were found using equations 17.100-17.114 from Raymer. It should be noted that the climb distance was reduced to zero because the calculations yielded that the obstacle was cleared during the transition to climb distance. These values are summarized in Table (9) below.

Stall Speed 85 knots Takeoff Speed 94 knots Approach Speed 128 knots Takeoff Distance 2600 ft Landing Distance 2400 ft

Table (9) Aircraft performance abilities

Structures The material chosen for the structures on the aircraft was a Carbon Fiber Reinforced Plastic (CFRP). The advantages that this material offered were that it has a high strength and is light weight. This material is predicted to give a 25% weight savings over existing aluminum aircrafts when 80% of the aircraft is composites by weight. This will allow team Arrival’s aircraft to weigh significantly less than modern day aircrafts of the same size. The increased strength of this material will also allow the aircraft to use forward swept wings without the fear of a structural divergence in the wings, which was a problem with past forward swept wing concepts. The biggest problem with using CFRP material is the cost. This material is currently expensive. Arrival is making the assumption that this material will become more common with future use and will decrease in price. If the price of this material were to decrease by the time the plan is commissioned in 2058 it would be the perfect material to meet the needs of our aircraft. A view of the structural layout of the aircraft wing can be seen in Figure (28) below. The wing contains two main spars. The ribs are spaced throughout the wing and are in the form of a supercritical airfoil which was previously discussed. The engine mount has been placed between the 4th and 5th rib from the wingbox and is used to hold the engine on the wing. The landing gear kick spar was added to its given location in order to aid in the placement of the center of gravity. The landing gear placed on the kick spar at its given location gave the most statically stable placement for the center of gravity. The wingbox carrythrough was incorporated to aid in the attachment of the wing to the fuselage.

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Figure (28) View of structural layout of aircraft wing

Weights and Balance The estimation of the weight of the aircraft is a critical part of the design process.

Using detailed statistical equations for various components will give more accurate weight values instead of just using crude component buildup based upon planform areas, wetted areas, and percents of gross weight. Using the areas calculated by the sizing code previously and equations 15.25 through 15.59, which are estimations for Cargo/Transport Aircrafts in Raymer’s book, a weight breakdown table was tabulated as shown below in Table (10). A weight savings of 25% over composites was applied to some components such as wings and fuselage. The matlab code used for these weight estimations can be found in Appendix (B).

Engine Mount Landing Gear Kick Spar

Wingbox Carrythrough

Main Wing Spars

Wing Ribs

31

Structures Weight (lbs)

Wing 5,911 Horizontal Tail 962

Vertical Tail 993 Fuselage 5,174

Main landing gear 2,554 Nose landing gear 357

Nacelles 2,456

Propulsion Weight (lbs)

Total Engine Installed 10,000 Engine Control 45

Starter 171 Fuel System / Tanks 196

Equipment Weight (lbs)

Flight Controls 1,169 APU Installed 2,323 Instruments 356 Hydraulics 299 Electrical 186 Avionics 124

Furnishings (w/o Seats) 502 Air Conditioning 934

Anti-ice 156 Handling Gear 23 Useful Load Weight (lbs)

Aircraft Empty 37,348 Crew (6) 1,200

Fuel 9,312 Passenger (177) 38,940

Table (10) Component weight breakdown

In addition, Figure (29) shows a broader weight breakdown view. Each part is a

percentage of the gross takeoff weight. Note that payload is a majority of the total gross takeoff weight.

32

Weight Breakdown as percentage of GTOW

System and Equipment

7%

Payload (crew+passenger)

48%

Propulsion12%

Structure22%

Fuel11%

Figure (29) Component weight breakdown

The average weight for each passenger with luggage was set at 220 pounds and

the average weight for each crew member with luggage was set at 200 pounds. With 177 passengers for single class configuration and 6 crew members, including pilots and flight attendants, were used in the calculation. By adding up the weights of the components, the aircraft’s manufacturing empty weight came out to be a total of 37,348 pounds. This empty weight correlates well with the results from the sizing code. Stability

After determining weights of each component, the next step was to figure out where the center of gravity was located at and what the static margin was, which will move with differing configurations. In order to achieve positive static margin for stability, weight components were forced to be shifted to various locations. The main factor was the wing location since it was a big portion of the total weight. Figure (30) illustrates where the major components are located in the aircraft.

33

Figure (30) Weight locations on aircraft

Center of gravity, neutral point and static margin are then calculated by Equations

(13) through (16).

.i

i iW xC G

W= ∑∑

(13)

iW = component weight

ix =distance from nose

. htac ht

w

aN P x C

a= + (14)

acx =wing aerodynamic center

htC =horizontal tail volume coeffiecient

hta =horizontal tail lift curve slope

wa =main wing lift curve slope

w

ht htht

l SC

MAC S=

i

(15)

htl =distance from center of gravity to tail quarter chord

htS =planform area of horizontal tail

wS =planform area of main wing

MAC =mean aerodynamic chord

. np cgx xS M

MAC

−= (16)

Wht

Wmaingear

Wengine

Wwing

Wnosegear

Wfuse

Wfuel

Wvt

Wapu

Wpayload

Wfurnishing Wcrew Wnacelle

34

npx =distance from nose to neutral point

cgx =distance from nose to center of gravity

Alternate payload configuration generates a different center of gravity location

and different static margins which can be seen below in Table (11).

Weight (lbs) C.G location Static Margin (%) Aircraft Empty 37348 65.77 45.9 Empty + Crew 38548 65.41 49.1

Empty + Crew + Fuel 47860 66.34 40.8 Takeoff 86800 69.83 9.7 Landing 78419 69.84 9.7

Table (11) C.G locations and S.M for different weights

Takeoff weight includes 177 passengers with luggage as well as 6 crew members

and maximum fuel. At takeoff, the center of gravity is at 69.83 ft from the nose while the main gear is located at 73 ft from the nose. Since the center of gravity is in front of the main gear, a tip over will not occur for this tricycle configuration aircraft. Landing weight is assumed to have only 10% of the total fuel left. Neutral point remains the same for all configurations which is located at 70.96 ft from the nose. Since fuel is placed in the wings, the center of gravity will not change a lot even when fuel is consumed during flight. Figure (31) demonstrates the static margin travel based on the numbers calculated.

Figure (31) S.M travel diagram

Fore C.G limit was 65.41 ft when only crew was added to the aircraft and the aft

C.G limit was 69.84 ft at landing. The static margin lies between 9.7% and 49.1%. Since OWE is not an actual flight condition, it can be said that the aircraft has about 10% of

Static Margin Travel Diagram

30000

40000

50000

60000

70000

80000

90000

5 10 15 20 25 30 35 40 45 50 55

Static Margin as % of MAC

Wei

ght (

lbs)

35

static margin for the whole flight. For any given condition, static margin is always positive so the aircraft is said to be statically stable. Tail and Control Surface Sizing

Vertical tail sizing was done by assuming one engine is out and the other engine had to produce the total amount of thrust to takeoff. This will generate a yawing moment about the C.G and in order to cancel out the moment, the rudder and the vertical tail have to be sized properly to generate enough lift. Since forward swept wings are used in the aircraft, the wing box tends to be located more aft than conventional aircrafts. Because of this, moment arms tend to be shorter than conventional aircrafts, thus increasing the vertical tail size. Using a symmetric airfoil such as NACA0012, lift coefficient can be calculated for various angles of attack. Figure (32) shows the free body diagram for the one engine out situation. The matlab code used for sizing can be found in Appendix (B).

Figure (32) One engine out free body diagram

Table (12) Vertical tail and rudder size

With a typical one hinge rudder used, the vertical tail had to be very large, so a

double hinge rudder was applied to provide more lift with a smaller area. The vertical tail

# of hinges Vertical tail size (ft^2)

Rudder size (ft^2) Area percentage deflection

(deg) 1 720 432 60 20

1 670 335 50 20

1 620 248 40 20

2 560 280 50 20

16ft

45.2ft

One Engine Thrust:

26900 lbs

Lvt

36

was eventually sized to be 560 ft2 with 50% double hinge rudder and 20 degrees of deflection to cancel out the yawing moment generated by one engine out case.

Horizontal tail sizing was done by calculating the moment about the main gear at 10 degrees of rotation. From the aircraft’s geometric configuration, the max rotation angle could go up to 14 degrees. However, to avoid tail strike for all circumstances, it was constrained to 10 degrees. The horizontal tail and elevator had to produce enough downward lift to takeoff from short runways. The same symmetric airfoil was used on the vertical tail as the horizontal tail. An airfoil such as the NACA 0012 was used to calculate various lift coefficients corresponding to different angles of attack. Figure (33) shows the free body diagram containing longitudinal forces. Again, the matlab code used for sizing can be found in Appendix (B).

Figure (33) Longitudinal forces free body diagram

By summing the moments about the main gear with 10 degrees of rotation angle,

Table (13) was tabulated and the horizontal tail was sized to be 450 ft2 with 50% elevator and 20 degrees of deflection to perform short takeoff.

Horizontal tail size (ft^2) Elevator size (ft^2) Area percentage deflection (deg)

410 246 60 20

450 225 50 20

480 192 40 20

Table (13) Horizontal tail and elevator size

Terminal Servicing On the advice of advisors at Boeing during the last presentation we created a terminal servicing diagram was created to show that the aircraft’s layout is serviceable at an airport. Figure (34) below shows the conventional servicing of the aircraft using equipment similar to what is used today. This diagram was based off of the terminal servicing guides posted by Boeing for the 737 series. Even though the service date for this aircraft is 2058, it is also designed to be able to serve smaller secondary airports.

Lht

µR

D

L

T

W Cm_wb

14º

37

Even at this date some airports may still be using older equipment similar to what is in use today, since 50 years ago similar equipment was servicing Boeing 707’s. The diagram in Figure (34) shows all equipment that would be necessary to service the plane without use of the APU.

Figure (34) Conventional Terminal Servicing Diagram

A more futuristic terminal servicing diagram was designed with the consideration that airports would modernize to fully take advantage of features of our aircraft. Figure (35) below shows the more futuristic servicing scenario which takes advantage of the hydrogen fuel cell APU on our aircraft. This configuration uses the hydrogen fuel cell APU to provide power for the aircraft, and air-conditioning system. By using a new hydrogen fuel truck, the APU can be run without draining the onboard hydrogen fuel supply. The new hydrogen truck eliminates the need for the electrical, pneumatic, and air-conditioning trucks, which helps eliminate extra equipment and simplifies servicing. It also greatly reduces the environmental impact of servicing our aircraft. Instead of power being produced by gasoline generator trucks or the jet engines which all produce harmful emissions, all of the power comes from the APU which only produces clean water as an emission. The water produced by the APU could be used to refill the aircrafts potable water supply which could potentially eliminate the need for a water truck.

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Figure (35) Futuristic Terminal Servicing Diagram

Environmental Impact The environmental impact of our aircraft has been one of the primary focuses in the design process. There are many different ways an aircraft impacts the environment. The main environmental impacts are noise output, and harmful emissions. Other considerations include the manufacturing waste in the production of the aircraft, and end of lifecycle disposal. The noise output of an aircraft is a major consideration, especially the noise at the airport during takeoff and landing. Advancements in design of jet engine nacelles and exhaust flow have helped quiet the noise of an engine. We looked at the improvements on the Boeing 787 that help with the reduction of noise to levels lower than every other aircraft. Arrival will make use of a chevron exhaust nozzles, and noise absorbing inlets like those on the 787’s nacelles. These improvements, combined with the upper surface mounted engines will help reduce and shield noise from the community around the airport. Aerodynamic noise from the aircraft as it moves through the air also contributes to the overall noise impact of the aircraft. The use of composite materials results in a smoother aircraft skin and reduced drag. The improved airflow from the composites and better aerodynamic design result in aircraft that have much less aerodynamic noise than their predecessors.

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Composite materials are not only better for noise impact, they’re better for manufacturing and disposal. Manufacturing an aircraft that is made largely from composites eliminates a large amount of production waste. When constructing an aircraft that is primarily aluminum, large volumes of metal are cutoff, scrapped, and must be melted and reformed before being used again. With composited the material is formed for the aircraft resulting in much less waste. Composites also eliminate the need for large amounts of rivets used to hold together an aluminum aircraft. Disposal of composite aircraft structures is a field that is currently being expanded and improved. With extensive use of composites in new aircraft like the Boeing 787 and Airbus A350, the market is being driven to provide better means of recycling composite materials instead of landfilling or incinerating them. The emissions of CO2 and NOx from an aircraft are the other major consideration for a new aircraft. Besides the engines, there is emissions which come from the APU, and all the ground servicing equipment powered by gasoline engines. By replacing the standard APU with a hydrogen fuel cell we were able to eliminate the harmful emissions that would have been produced by a conventional APU as well as reduce the emissions from ground servicing vehicles.

Maintenance and Reliability Ease of maintenance is important for reducing operating cost for airlines and aircraft downtime. The choice to use composites for a majority of the aircrafts construction was made to reduce weight, but it also reduces maintenance. Composites typically do not experience fatigue to the same extent that aluminum does. Furthermore, composites do not corrode. The engine nacelles of our aircraft have also been designed for ease of maintenance. The design of the nacelles was based off that of the Boeing YC-14, which had nacelles that focused on engine serviceability. The low wing allows for easier access to the nacelle versus the high wing of the YC-14. Access to the underside of the engine nacelle requires only a small platform for a mechanic to gain access to the engine core and components inside. Access to the rear of the engine is easily achieved because a mechanic can walk onto the wing and up to the engine. The use of fly by light fiber optics to replace fly by wire copper wires will also decrease aircraft maintenance. Fiber optic wire is much more durable and does not suffer from corrosion like copper. The use of advanced electronic avionics along with the fiber optics reduces mechanical complexity, and increases the reliability of the flight systems. The cabin will be lit by LEDs instead of incandescent bulbs like those used on the 787. LEDs have a much longer lifespan than traditional light bulbs which dramatically reduces maintenance and part costs over the life of the aircraft. By 2058, much more durable and longer lasting materials will be available for use in the passenger cabin thereby increasing the lifespan of seats, tray tables, overhead bins, and many more parts worn down by continual use.

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Summary Team Arrival is confident in the aircraft concept we have created. It has successfully met all of the criteria we initially laid out and incorporates many new and advanced technologies. Each feature added to the aircraft has served to make it a better, safer, more eco-friendly aircraft than the ones that have come before it. We feel that the aircraft is properly configured to meet the servicing needs of several different types of airports because we are targeting many different airports for our missions. We also feel the performance characteristics of the aircraft will make it appealing to airlines and the comfort of travel will make it a top choice for their passengers.

Our initial plan was to create an aircraft capable of taking off and landing in extremely short distances in order to utilize secondary airports as well as half-runways at larger hubs. This was to relieve the continuously growing congestion that is mounting at airports across the country. Team Arrival is confident that, with the incorporation of our aircraft, airports will be able to better manage the traffic they are experiencing. We have taken great pride in providing this aircraft concept and feel that it will meet all of our initial customer’s needs and then some.

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Appendix A Aircraft Dimensions

18.5 102 ft

123 ft 63 10

35

14

42

16 ft

40 ft

Fuselage Width:

12ft 4in

Wing Area:

1032 sq ft

H. Tail Area:

450 sq ft

43

Bibliography i Daniel P. Raymer, Aircraft Design: A Conceptual Approach, 4th ed. (Location: Publisher, Year) Pages-Pages. ii “Pratt & Whitney hopes ride on geared fan”, Flug Revue, Feb. 2007, 22 April 2008, <http://www.flug-revue.rotor.com> iii “Press Release”, Pratt & Whitney, 18 July 2006, <http://www.pratt-whitney.com> iv Philip G. Hill, and C. R. Peterson, Mechanics and Thermodynamics of Propulsion, (Location: Publisher, Year) Pages-Pages. v Englar, Robert J., and Nichols Jr., James H., “Advanced Circulation Control Wing System for Navy STOL Aircraft”, AIAA Paper 80-1825R, Dec. 1981