Development Process/General Areas (MDJ)

AAE 451

Systems Definition Review

Group 6

John Collins Chad Davis Chris Fles Danny Sze Ling Lim Justin Rohde Ryan Schulz Ronald Wong Yusaku Yamashita Contents

Executive Summary...... 3 1. Review of Business Case...... 4 2. Current Design Requirements...... 5 3. Trade Studies...... 6 4. Concept Generation (Pugh’s Method)...... 19 5. Selected Concepts...... 23 6. Preliminary Layout...... 27 7. Fuel Choice...... 33 8. Constraint Models...... 35 9. Aircraft Characteristics and Comparisons...... 37 10. Conclusion...... 37 11. Appendix...... 38 12. References...... 69 13. Merit Pool...... 70 Executive Summary

The current outlook on the world’s fuel supply shows the possibility of nearing peak oil within the next few decades. This situation will drive up the cost of fuel to possibly unattainable prices. The occurrence of peak oil, however, will not eliminate the need for individuals to travel and for business to work fluidly. This report assesses the feasibility of creating an aircraft that is capable of safely and efficiently operating on an alternative fuel.

The design team believes that a market exists for a small to medium sized propeller driven aircraft. The market to which this could be sold to consists primarily of air taxi services, air charter operations, and corporate flight departments. The concept aircraft could also be sold for uses in cargo, medical, and evacuation roles. The primary markets exist not only in the United States, but also in Europe and Asia.

The team feel that the best way to effectively capture a portion of these markets is through the development of a small single turboprop aircraft. It will be capable of carrying six passengers and two crew members on a 600 nautical mile trip. With the capability of using runways as short as 2,100 feet, the aircraft will be able to utilize a multitude of airports providing more convenient point to point service.

A main focus of this design will be to maintain affordability and low cost of the aircraft. A lower acquisition cost and operating cost will positively affect both the operator and end-user of the aircraft, which will in turn bolster sales. Through the use of trade studies and concept generation and selection process (Pugh’s method), a design has been chosen to suit these criteria. This report also outlines potential interior layout, as well as alternative fuels available for our propulsion requirements.

The combination of an alternative fuel aircraft, a new and broad market, and effective utilization of the many advanced design tools available today will make the aircraft a successful competitor. 1. Review of Business Case

Energy forecasts provided by Energy Information Administration (EIA) and ExxonMobil reflect the best case scenario for future petroleum prices. Figure below demonstrates that with external factors, the price trend is likely to rise for the next 30 years:

Fuel Price Forecast

) 200 n o l l a g / s t

n 150 e c (

e c i r

0 1985 1990 1996 2001 2007 2012 2018 2023 2029 Year

Historical Jet Fuel Price Projected Jet Fuel Price (EIA) Projected Ethanol Fuel Price Projected Jet Fuel Price (based on historical data)

The current aim is to enter the market around the year 2018, when the alternative fuels become competitive.

Overall, the acquisition and operating costs of the alternative fuel based aircraft need to be very strict in order to compete in the current market. Currently, market data shows direct operating costs for light turboprops are around $500-$600 per hour, while small jet aircraft cost approximately $1,500-$1800 per hour. However, this includes aging aircraft such as the KingAir C90s designed in the 1970s, and with the advancement in technology, these figures are likely to decrease. In terms of acquisition costs, the alternative fuel aircraft would need a purchase price of about $1.6M if it was propeller driven aircraft or about $1.8M if it is a jet aircraft. 2. Current Design Requirements

The current design requirements are:

Cabin Capacity 2 crew + 6 passengers Cruise Range (nm) 600 Cruise Speed (kts) > 250 T/O distance (ft) < 2100 Acquisition Cost ($M) < 1.8 DOC ($/hr) < 550

Table 1 - Current Design requirements

These values were obtained from the QFD matrix and verified using the trade studies discussed in section 4. They represent the limits on which further design processes will be based.

Figure 1 below shows a typical design mission profile:

Figure 1 - Design Mission Profile 3. Trade Studies

Trade Studies:

The trade studies conducted for this report had two major impacts on the design process. First, the trade studies provided justification for the design requirements that any conceptual designs need to meet, as well as improving the accuracy of the design requirements that were derived from the QFD Matrix. The second major effect of the trade studies was the justification of the ratings made during Pugh’s Method in section 4.

Initial Sizing:

In order to conduct the trade studies, an initial sizing code needed to be created to determine the aircraft weight based off of characteristics such as L/Dmax, specific fuel consumption, and empty weight fraction. The sizing code uses the aircraft characteristics and the design mission profile to calculate weight fractions for different segments of the flight based off of an initial guess, and then returns the calculated aircraft weight using Equation 1 below. The calculated value then becomes the initial guess in an iterative process to determine the actual aircraft gross weight.

w  w w  crew payload 0 w w Equation 1 1- fuel - e w 0 w 0

The aircraft characteristics such as specific fuel consumption and L/Dmax are determined using estimates from historical data of aircraft of the same size and type, as well as from the competing aircraft database.

The empty weight fraction is determined by using a linear regression of the historical empty weight fractions for aircraft of the same size and type. The empty weight fractions for light business jets, single turboprop aircraft, and twin turboprop aircraft were determined as a function of gross takeoff weight using Equation 2, where A and B are constants determined from the linear regression. The constants for the single turboprop aircraft and the light turbofan aircraft are shown in the MATLAB Code Appendix.

w e B  A  w 0 Equation 2 w 0 The fuel fraction is the function of each of the weight fractions of the design mission segments. Many of the weight fractions are taken from historical values such as warmup, takeoff, and landing. Other mission segments, such as climb, cruise, and loiter, have weight fractions that are entirely dependent on the mission profile. Because turbofan and turboprop aircraft have very different climb, cruise, and loiter performance and conditions, it is important to accurately calculate these weight fractions in order to better distinguish between these two classes of aircraft for Pugh’s Method. The climb weight fraction is based on Mach number. Turbofan aircraft climb to higher altitudes and higher Mach numbers than turboprop aircraft, so the turbofan aircraft will burn more fuel as can be seen in Equation 3. The cruise and loiter weight fractions are based on the Breguet range equation and the endurance equation and are shown in Equation 4 and Equation 5, respectively. Turbofan and turboprop aircraft also fly at different L/D values, and this is accounted for in the sizing code. Once the mission segment weight fractions are calculated, the fuel fraction is determined using Equation 6, which makes an allowance for trapped fuel.

w i  1.0065 - 0.0325 M Equation 3 w i-1

w i  - RC   exp  Equation 4 w i-1  VL/D

w i  - EC   exp  Equation 5 w i-1  L/D 

w fuel  w x   1.061-  Equation 6 w 0  w 0 

Once the aircraft gross weight is calculated, further studies can be conducted. Aircraft acquisition and direct operating costs (DOC’s) can be determined with a linear regression similar to the empty weight fraction. Equation 7 shows the cost model, which is based on aircraft gross weight, range, and speed. The constants A, B, C, and D are based on the linear regression for each aircraft type, and are different for the acquisition costs and DOC’s. The values for the cost regressions are shown in the MATLAB Code Appendix.

A B C Cost  w 0  R  V  expD Equation 7

Trade Study Analysis:

The results of the trade studies can be seen in their entirety in the Trade Study Appendix.

The trade studies show clear trends in aircraft gross weight, acquisition cost, and DOC as values such as specific fuel consumption, range, and speed, are varied. The results of the trade studies were critical in determining whether turbofan or turboprop aircraft were the best choice for our mission, and certain design criteria were also created from the studies as well.

When the range of the mission was varied, the aircraft gross weight increased significantly, while the DOC per hour increased slightly for turbofan aircraft, and almost not at all for turboprop aircraft. The trend for the acquisition cost shows a decrease in cost as range increases, but this is entirely based on the linear regression of similar aircraft, and in the case of very light jets (VLJ’s), there are only a few examples so the data fidelity is in question. The true acquisition cost probably follows the gross weight calculation, which is affected by the mission profile and not the cost regression.

The trends for acquisition cost, DOC per hour, and gross weight for turbofan and turboprop aircraft are exactly opposite. The turboprop aircraft had lower costs and weights as the speed decreased while the turbofan aircraft had lower DOC’s per hour, gross weight, and thus acquisition costs, as the cruise speed was increased. If the turboprop aircraft operates at 250 knots and the turbofan aircraft operates at 350 knots, the turbofan aircraft will save the passengers 30 minutes after a flight of 600 nautical miles. This shows that while the costs may decrease for slower flying turboprop aircraft, customers expect a timely arrival, and flying slower than 250 knots will give a significant edge to turbofan powered aircraft. As the speed of the turbofan aircraft increased, the costs decreased, but the cost reduction was diminishing. This is most likely due to the fact that turbofan aircraft have higher optimal cruise speeds than turboprops. Flying faster is more cost effective for turbofan aircraft until the Mach number reaches a certain point where the increased drag causes more problems. Turbofan aircraft of our size typically operate at 350 knots and an altitude of 30,000 feet. Figure 2 and Figure 3 show the gross weight trends for range and speed. J e t - T r a d e S t u d y ( R a n g e a n d S p e e d ) 1 3 5 0 0

1 3 0 0 0 2 0 0 k t s

1 2 5 0 0

2 2 5 k t s 1 2 0 0 0 ) s b l 1 1 5 0 0 2 5 0 k t s ( t h g i 2 7 5 k t s e 1 1 0 0 0 W

s 3 0 0 k t s s 1 0 5 0 0 o r 3 2 5 k t s G 1 0 0 0 0 3 5 0 k t s

9 5 0 0

9 0 0 0

8 5 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 2 - Turbofan Range and Speed Trade Study

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 7 5 0 0 3 5 0 k t s 3 2 5 k t s 3 0 0 k t s 2 7 5 k t s 2 5 0 k t s 7 0 0 0 2 2 5 k t s 2 0 0 k t s ) s b l 6 5 0 0 ( t h g i e W s s 6 0 0 0 o r G

5 5 0 0

5 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 3 - Turboprop Range and Speed Trade Study At the time the study was conducted, the exact fuel type and engine had not been selected, so the specific fuel consumption was varied to create design requirements for the fuel type and engine. For both the turboprop and turbofan aircraft, a specific fuel consumption greater than 0.7 lbf/(lbm∙hr) resulted in aircraft gross weights, acquisition costs, and DOC’s that were too high to remain competitive in the current market, as can be seen in Figure 4 and Figure 5. Similarly, aircraft with an L/Dmax ratio less than 12 could not remain competitive, which may show that aircraft with canards will have an advantage over conventional aircraft.

J e t - T r a d e S t u d y ( R a n g e a n d S F C ) 7 0 0

0 . 8 S F C 6 5 0

6 0 0 0 . 7 S F C r u o H /

s 5 5 0 r 0 . 6 S F C a l l o D

5 0 0 0 . 5 S F C

4 5 0 0 . 4 S F C

0 . 3 S F C 4 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m ) Figure 4 - Turbofan Range and SFC Trade Study T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 3 6 0 0.8 SFC

0.7 SFC

3 4 0 r u

o 0.6 SFC H /

s 3 3 0 r a l l o D 0.5 SFC 3 2 0

0.4 SFC 3 1 0

0.3 SFC 3 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 5 - Turboprop Range and SFC Trade Study

Overall, significant performance issues between turboprop and turbofan aircraft were observed. Twin turboprop aircraft had the highest gross weights, acquisition costs, and DOC’s. Based on the trade studies, single turboprop aircraft are significantly more cost efficient and have similar performance capabilities when compared to twin turboprop and turbofan aircraft. From the trade study analysis, the single turboprop aircraft is optimal for the current design mission. However, the public perception of propeller driven aircraft is not based on efficiency, and factors other than efficiency and performance also determine which aircraft best suites the customers needs. This is discussed further in Pugh’s Method.

FLOPS optimization software was used in conjunction with the regression models to investigate the variation of operating cost with the cruise altitude at various cruise speeds using existing aircraft database. Three aircraft with different power plants were used to initiate the FLOPS software. The King Air C90 was used model the twin turboprops; the Pilatus PC-12 was used to model single turboprops and the Eclipse 500 was used to model the very light jets. By varying the cruise altitudes and cruise speeds in the FLOPS input file, the operating costs were obtained and plotted using excel. As there were many uncertainties with the input into FLOPS, only the trends predicted by FLOPS of the operating costs were noted as the exact values were highly inaccurate. A 5% uncertainty error in the operating costs was also introduced into the very light jets model as a safeguard against overly cautious performance data obtained from the manufacturers’ website.

TWIN TURBOPROP

Operating Cost versus Speed t s o C

g n i t a r e p O

0.3 0.35 0.4 0.45 0.5 Speed (Mach)

500nm 550nm 600nm 650nm 700nm 750nm 800nm

Figure 6 - Operating cost versus cruise speeds at 10,000 ft over various ranges

 It can be observed that as the range increases the operating costs also increased.  The operating cost decreases as speed increases and is at the lowest at mach 0.5 (320 knots).

Operating Cost versus Different Cruising Altitude

t s o C

g n i t a r e p O

5000 10000 15000 20000 25000 30000 35000 Height(ft)

500nm 550nm 600nm 650nm 700nm 750nm 800nm Figure 7 - Operating cost versus cruise altitude at mach 0.45 over various ranges t s o C

g n i t a r e p O

5000 10000 15000 20000 25000 30000 35000 Height(ft)

500nm 550nm 600nm 650nm 700nm 750nm 800nm Figure 8 - Operating cost versus cruise altitude at mach 0.5 over various ranges  It can be observed that as the range increases the operating costs also increased.  Above 10,000 feet the variation of operating cost with altitude is minimal. However acquisition cost might increase at higher altitude due to pressurization requirements.  Cruise Altitude ~ 10,000 feet, Cruise Speed ~ 320 Knots

SINGLE TURBOPROP

Operating Cost versus Different Cruising Altitude s t s o C

g n i t a r e p O

5000 10000 15000 20000 25000 30000 35000 Height (ft)

500nm 550nm 600nm 650nm 700nm 750nm 800nm

Figure 9 - Operating cost versus cruise altitude at mach 0.3 over various ranges t s o C

g n i t a r e p O

5000 10000 15000 20000 25000 30000 35000 Height (ft)

500nm 550nm 600nm 650nm 700nm 750nm 800nm Figure 10 - Operating cost versus cruise altitude at mach 0.4 over various ranges

 It can be observed again that as the range increases the operating costs also increased.  The variation of operating cost with altitude is minimal. However acquisition cost might increase at higher altitude due to pressurization requirements.  Cruise Altitude ~ 10,000 feet, Cruise Speed ~ 255 Knots VERY LIGHT JETS

Operating Cost versus Speed t s o C

g n i t a r e p O

0.51 0.52 0.53 0.54 0.55 Speed (Mach)

500nm 550nm 600nm 650nm 700nm 750nm 800nm Figure 11 - Operating cost versus cruise speeds at 30,000 ft over various ranges

 It can be observed that as the range increases the operating costs also increased.  The operating cost decreases as speed increases and is at the lowest at mach 0.55 (325 knots). Operating Cost versus Different Cruising Altitude t s o C

g n i t a r e p O

24000 26000 28000 30000 32000 34000 36000 Height

500nm 550nm 600nm 650nm 700nm 750nm 800nm Figure 12 - Operating cost versus cruise altitude at mach 0.51 over various ranges t s o C

g n i t a r e p O

24000 26000 28000 30000 32000 34000 36000 Height(ft)

500nm 550nm 600nm 650nm 700nm 750nm 800nm Figure 13. Operating cost versus cruise altitude at mach 0.55 over various ranges

 It can be observed again that as the range increases the operating costs also increased.  30,000 feet is the optimum cruise altitude for the very light jet model.  Cruise Altitude ~ 30,000 feet, Cruise Speed ~ 325 Knots

SUMMARY t s o C

g n i t a r e p O

500 550 600 650 700 750 800 Range

Very Light Jet Single Turboprop Twin Turboprop Figure 14 - Comparison of Operating cost for the Different Power Plants

 Operating cost of twin turboprop is the highest regardless of the range.  Operating cost for the single turboprop is the least regardless of the range. 4. Concept Generation

Pugh’s Method

The aim of Pugh’s method is to select the best concept design for our market sector. The approach used was to first gather potential designs by brainstorming. These were then compared against important criteria that would define a successful product. Since the future of air charter/air taxi/corporate flight departments require an affordable means of travel, criteria are grouped under the categories of acquisition and operating costs.

The individual characteristics of each concept is rated as either positive (+), negative (-), or same (s) in comparison to the benchmark aircraft (Beechcraft King Air C90). The design mission requirements define the lower limits set for each criterion. A concept can better these limits if they operate at a higher efficiency. For example, given the airspeed is set at 250 kts, a jet variant can fly at 350 kts if it is consuming less fuel at that speed.

Table 2 – Pugh’s Method

, a a ) ) t t

) n k , l e h o c i l j e i i ( g a a t n i t a i b t n ( d h

e r e r , d t n a v i e g f e n s n

( n i m u

r o i l

s i , o e c

w g b g h

3rd iteration - concepts to be carried forward ,

m a r n s n

d t i i e i u e s l

l w p w g l m

u , a a e , l s p l w c s

e i d g o e t s e r l r l n n u a i i e g f n

S g v n a 2 i n C S e 15 1 13 Desirable Characteristics/Criteria op cost Range s s s acq cost Compatibility with T/O and landing (pitch) s s s op cost Performance SFC s s - acq cost Stability s s s acq cost Flyability s s s acq cost Manufacturability s s s acq cost Noise s - s acq cost Emissions s s - acq cost Cabin Comfort s s - acq cost Pressurisation of cabin s s + op cost Maintainability s s - op cost Reliability/Safety s - s acq cost Development Cost s + + Public Perception s + + Concept 15 Total Canards, wings (far back), vertical + 0 2 3 stabiliser (tail), single pusher S 14 10 7 engine - 0 2 4 Each criterion is explained below. Data for deciding the efficiencies are obtained from trade studies done in previous sections.

Factors Affecting Operating Cost

Range – efficiency of structure and engines (measured by drag and SFC) over given distance

- For a given range, singles burn less fuel than twins for same speed (250kts) - Turbofans are slightly better than twin turboprops, and are comparable to single turboprops. - Canards give less drag which will improve SFC, thus cost.

SFC – consumption rate of engines

- single engines cost less than twins for a given SFC - turbofans are comparable to single turboprops for given SFC - For a SFC increase from 0.5 to 0.6: Single turboprops increase in cost by +$20/hr; twin turbofans increase by +$50/hr; twin turboprops by +$100/hr

Maintainability – ease of servicing and turn around time

- The number of engines is directly proportional to the cost of maintenance - high tails and boom tails are more harder to maintain

Factors Affecting Acquisition Cost

Compatibility with T/O and landing – pitch ease/clearance of tail and propellers at T/O and landing

- pusher-engines are less able to clear the ground than tractor ones

Reliability/Safety – glide ability in the event of an engine out

- canards are naturally less stable without active controls, but offer less drag and better lift to glide than conventional configuration - twin engines have bad yawing characteristics with engine out - blended wing body will have less corrective-yaw control in case of engine out Stability – stall characteristics

- better stall characteristics as it offers no stall over main wing with the canard itself stalling earlier - boom tail have worse stall characteristics - whole blended wing may stall catastrophically

Flyability – conventionality and placement of control surfaces with regards to pilot's view

- unconventional designs such as boom tail, far back wing, canted canards lead to commonality problems, increasing pilot training costs, and handling issues

Manufacturability – production cost, counting from frozen design specs and design

- Boom tails, 3-engine configuration, coaxial propeller design, blended wing body, canted-canards, high tail designs are all more expensive to manufacture than conventional configuration.

Noise – internal cabin passenger noise

- Pusher-engines are less noisy - Turbofans are less noisy than propeller-driven engines - Fuselage mounted configurations have worse vibration problems than wing mounted ones

Emissions – emissions of greenhouse gases, NOx and other particulates

- Single engines are more efficient, thus burn less fuel than turbofans and twin engines (SFC’s), leading to lower emissions. i.e. for a given emission requirement, single engines will cost less than twins and turbofan engines based on their SFC values

Cabin Comfort – meeting the market cabin size (an initial estimate of 5’ x 5’)

- Turbofans have smaller cabin sizes than turboprops: Turbofans: Eclipse 500 – 4.2’ x 4.3’i Citation Mustang – 4.6’ x 4.7’ii Turboprops: Pilatus PC-12 – 4.9’ x 5.0’iii

Pressurization of Cabin – ease of incorporating a pressurised cabin, system redundancies in meeting FAR requirements

- Single engines will require a redundant system twin engines already have dual feed/source system for pressurisation - Jets will require a pressurised cabin due to their cruise altitude Development Cost – cost of developing new systems such as the avionics needed to control the canards

- radical and unconventional design such as canards, boom tails and blended wing body will lead to higher development costs

Initially, 15 concepts were drawn up by the team and weighed against the same criteria. (See Table 4 for the 1st iteration). From the resultant ratings, concepts 1, 4, 8, 13 and 15 were carried through to the next iteration, with concept 15 used as the new benchmark. Concepts 4 and 15 were very similar, and therefore one was eliminated.

More concepts were thrown into the second iteration (See Table 5 for the 2nd iteration). These were compared with concept 15, and the final three concepts (concepts 1, 13 and 15) were selected and compared. (See Table 2 for the 3rd iteration). 5. Selected Concepts

(a) Concept 1 (b) Concept 13 (c) Concept 15

Figure 15 – Selected Concepts

Concept 1 features a low wing, single tractor turboprop engine design, with a conventional mid horizontal tail and vertical stabiliser, as well as a tricycle landing gear. It was chosen due to the simplicity in its proven design and conventional configuration, leading to a low development, high reliability and good manufacturing costs.

Concept 13 features a low wing with winglets, twin fuselage-mounted turbofan configuration, and a high tail design, again with a tricycle landing gear. This was chosen to allow flexibility in our design by including a jet variant.

Concept 15 features a low wing placed at the aft-end of the fuselage, with a canard configuration, vertical stabiliser and a single pusher turboprop engine, complete with a tricycle landing gear, chosen for its low drag and good stall characteristics, as well as lower noise than the conventional tractor variant. Table 3 summarizes the predicted performance for these concepts. Concept 1 Concept 13 Concept 15

GTOW 6800 lb 10000 lb 6800 lb W/S (wing loading) 32 lb/ft2 40 lb/ft2 32 lb/ft2 S (wetted wing area) 212 ft2 250 ft2 212 ft2 AR 7.6 8 7.6 b (wing span) 40 ft 1 in 44 ft 7 in 40 ft 1 in P/W (power to weight) - props T/W (thrust to weight) – P/W = 0.088 T/W = 0.155 P/W = 0.088 turbofans/jets

STO 2100 ft 2100 ft 2100 ft

VCruise 250 kts 350 kts 250 kts $1.85 $1.65 million $1.65 million Acquisition cost million 100LL - $350/hr 100LL - $350/hr $550/hr Direct Operating cost Biodiesel - $468/hr Biodiesel - $468/hr

Both acquisition and Table 3 - Concept Predicted Performance Values direct operating costs are calculated using regression models obtained from the aircraft database. (See appendix for the MATLAB code and graphs)

Biodiesel fuel is projected to burn 20% more than traditional 100LL fuel, thus with the assumption of a 20% additional pricing when compared to current fuels, a 44% increase in direct operating cost is added to give $468 for the turboprops.

Upon further discussion, turbofans are rejected based on the fact that the DOC costs are higher than the single turboprop designs and the wing loading are higher than required. Therefore, only concept 1 and concept 15 will be considered in the next phase. Aircraft Design Group 6

Table 4 - Pugh's Method, 1st Iteration 26 Aircraft Design Group 6

Table 5 - Pugh's Method, 2nd Iteration

27 Aircraft Design Group 6

6. Preliminary Layout

Fuselage Design

The operational characteristics of our aircraft are; - Standard 6 passengers - Canard - Single Turboprop - Pusher - Lower Dihedral wing (Tail configuration is not specified yet)

Figure 16 – Group VI Preliminary Fuselage Design

Why Canard?

The typical aircraft configuration includes all control surfaces aft the main wing. The difference between the horizontal control surface being located forward or aft of the main wing lies in the relative area of the control surface, as well as the location of the center of gravity of the aircraft.

Advantages of Canard

- Stall of the main wing may be virtually eliminated with careful design. The properly designed canard will stall before the main wing, rotating the nose downward and restoring lift as the velocity increases. - Both the fore and aft surfaces provide lift in the same sense. When properly trimmed they are capable of higher lift-to-drag ratios, having superior range compared with conventional aircraft. The foreplane lift is additive, in the same direction as the wing, helping to rotate the aircraft into the attitude required for lift off, unlike a conventional machine with the tail plane and elevator at the back.

Disadvantages of Canard

- The rear engine location of pusher airplanes is somewhat less desirable than a forward location. The propeller operates in a more turbulent airflow (caused by airflow over the wings), which decreases propeller efficiency. - The main wing becomes less efficient because it operates in the wake of the canard, thus lowering the available maximum lift. - Sizing of the canard is more difficult and has a larger effect on the stability of the aircraft than a tail configuration.

Dimension

Figure 17 - External Fuselage Dimension

The preliminary dimensions of the six-passenger single turboprop pusher aircraft are shown in Figure 17 above. The numbers are approximate, and may vary depending on engine size. Winglets may also be utilized in the final design, eliminating the need to a vertical tail.

Group VI Aircraft Single Turbo Prop Avg. Total Length 35-40ft Avg. Total Length 34.7 ft Wing Span 40 ft Avg. Wing Span 43.55 ft Overall Height 17 ft Avg. Overall Height N/A Tail Height 12 ft Avg. Tail Height 12.5 ft Fuselage Height 5 ft Avg. Fuselage Height 4.8 ft Wing Area 212 ft2 Avg. Wing Area 228.4 ft2 Cabin Volume 330 ft3 Avg. Cabin Volume 258 ft3 Table 6: Comparison of Group VI Aircraft and Average Single Turboprop Dimensions

As in Table 6 above, the aircraft dimensions are very similar to most other six-to-eight passenger single turboprop aircraft. The main advantage the aircraft has over competing models is the significant increase in cabin volume – an important factor in passenger comfort. However, our aircraft successfully minimizes the total wing dimensions. Therefore, total drag acting on aircraft is the same.

Nose and Tail Dimensions

The fuselage cone is normally a smooth transition from the maximum fuselage cross section to the “end of the fuselage. When the ‘fineness ratio’ of this cone is too low, there will be a large base drag penalty although the fuselage weight may be reduced. When the ‘fineness ratio’ of this cone is too large, there will be a large friction drag penalty as well as a large weight penalty. It will be obvious that a long fuselage cone tends to increase the tail moment arm thereby reducing required tail area and vice versa. The decision on the fuselage cone fineness ratio is given as below.

For Nose Cone Ln 1.5to2.0 D Eq .1

For Tail Cone Lt  2.5to3.0 D Eq. 2

Based on equations above efficient nose and tail length can be calculated. For our case, since our target D (height of Fuselage) is 5 feet tall, nose cone length should be approximately 10 feet, and Tail cone length should be approximately 15 feet. For canard pusher aircraft case, the engine of the aircraft has to be in the tail section, so that that tail length should be little longer and wider than usual.

Human Factors

For many years the weight of the average man was 170lb (77kg); the western Caucasian male is heavier and bulkier than his pre-World War II counterpart. Racial variations in human size and shape can make a considerable difference to weight and balance calculations and to fuselage proportions. The average weight today is nearer 180lb to 200lb clothed, with deviation of plus or minus 14%. Measurements also vary by plus or minus 5%.

The bulkier the pilot or passenger, the greater the attention must be paid to legroom and space for movement. This should include movement while seated, and the ability to interact with other passengers during flight.

Figure 19 - Seat Configuration

Above is the seat configuration of our aircraft. The minimum width of fuselage of fuselage for a single seat (or tandem) airplane is 24-in (0.61m). A side by side arrangement for large occupants must be at least 40-in internally (1.02m), while 42-in (1.1m) is common. The widest and most comfortable cockpit was 47-in (1.2m) internally. The narrowest is 38-in with a mere hand-hitting distance from seat to roof of 34-in (0.87m). If an airplane has a cabin with an aisle between the seats, aisle width should not be less than 6-in (our aircraft aisle width is around 20 to 24-in).

Wing Position

Low-winged airplanes provide a better arrangement when agility is required, because they tend to have good fields of view in the direction of turn and manoeuvre. The wings blanket downward view somewhat in straight and level flight. They generate more favourable ground effect on take off and landing. Low wing structure is also useful anchorage and stowage for landing gear, which can be made shorter and lighter. Deeper spars can often be used, by incorporating them into seat structure, without increasing the depth of fuselage needed for a high winged airplane with a wing of the same thickness.

Cabin Design

Selection of cabin cross section is critical to passenger comfort and to weight and wetted area. The fuselage cross section in Figure 19 is a typical “circular” cabin section which is usually good for pressurization. Because of the human anatomy it will be discovered that the fuselage will become rather bulky. This is the reason why the fuselage cross section of smaller general aviation airplanes is more or less rectangular.

Figure 19 – Cross-sectional View of Cabin Configuration

Figure 20 - Top View of Cabin Configuration

Typical spacing of each seat is around 34” to 36” for passenger aircraft. As shown in Figure 5, 40” of spacing should provide passengers more legroom and comfort during flight. The thickness of the fuselage is typically 1 to 4 in. This depends on whether the aircraft is pressurized or not.

Cabin Variation

Standard

- 2 pilots - 6 passengers - Baggage space in between cockpit and cabin

Maximum: 8 people (w/ baggage)

Executive

- 2 Pilots - 4 Executive Passengers - Baggage compartment in between cockpit and cabin.

Maximum: 6 People (w/ baggage)

Economy

- 1 Pilot - 8 passengers - No Baggage compartment

Maximum: 9 people (w/o Baggage)

7. Fuel Choice

Alternative Fuels

The group examined the properties of several different alternative fuels for use in twin piston, single turboprop, and twin turbofan driven aircraft. The advantages and disadvantages of the best choices will be discussed below.

Ethanol Blend - AGE85

AGE85 (Aviation Grade Ethanol) is a well established and tested ethanol based alternative fuel. It is specifically blended for cold starting, which makes it an ideal fuel for use in aviation where fuel line and carburetor icing must be avoided. It also burns much cleaner than traditional aviation fuels.

The major disadvantage of AGE85 and other ethanol-based fuels is its dangerous effects on fuel system components. Ethanol may react with seals or lines, causing corrosion. Significant modification of these parts would be necessary if an ethanol based fuel were to be used. Furthermore, ethanol has a lower energy density than traditional aviation based fuels. Lastly, the research with AGE85 has been primary focused on driving piston aircraft performance.

Pure Biodiesel – B100

Biodiesel is made mostly from soybean oils, and contains no petroleum products. Because of this, availability and affordability of biodiesel would not be directly affected by the world-wide petroleum markets. This is an important economic advantage in a situation where petroleum becomes prohibitively expensive. B100 already has a significant production infrastructure, so the availability of the fuel would not be in question. Other advantages of biodiesel include the fact that diesel reduces engine wear, is comparatively safe to store and transport, and has some of the lowest harmful emissions of any of the alternative fuels studied.

There are two major disadvantages for B100. The first is its high freezing and cloud points, as seen in Table 6. This causes problems for both operation and shipping of the fuel. Electric heaters would have to be used on the fuel tanks and engines during storage of the aircraft, and the shipping costs of the fuel would be significantly higher in cold- weather climates. The other major disadvantage is the reduced energy content. As Table 7 shows, the energy density of B100 is almost 18% lower than it is for traditional aviation fuels. Beyond the high freezing point and clouding point, researchers have discovered a problem with using soybean-based biodiesel due to the limitations in growing soybeans fast enough to produce the required fuel needed for the number of flights required.

No. 2 Diesel (petrol) B100 (pure biodiesel) Cloud Point (F) -9 35 Pour Point (F) -17 32 Table 6 – Temperature Characteristics of Alternative Fuels

Jet-A Avgas No. 2 Diesel B100 (pure (petrol) biodiesel) Heat of Combustion (Btu/gal) 123099 115480 131295 117093 Density (lb/gal) 6.676 6.092 7.079 7.328 energy density by mass 18439 18956 18547 15979 (Btu/lb) Table 7 – Comparison of Alternative Fuels to Existing Fuels

Biodiesel Blend – B20 & others

B20 and other like fuels are blends of petroleum-based diesel fuels with pure biodiesel. Blending biodiesel with petroleum-based diesel relieves some of the problems with pure biodiesel. The freezing point is lower, and the energy density is higher than biodiesel making it a more feasible choice.

However, the production of B20 still involves mostly petroleum. Therefore, B20 can not be considered a renewable fuel, and the price of the fuel would change with the petroleum market, which would be a major disadvantage when petroleum is expensive.

BioJet Fuel

The University of South Dakota is currently developing a bio-matter based fuel that has very desirable properties. Freezing is not a problem for this fuel, as it operates normally down to -75 F. It has very low emissions, and little modification would be needed to run in current turbofan engines. This BioJet fuel would be the ideal choice for an alternatively fueled aircraft.

However, this fuel is only in the experimental phase, and no infrastructure for the production and distribution exists. Designing an aircraft around these obstacles could be a large risk.

Figure 21 summarizes the advantages and disadvantages of each fuel type.

Figure 21 - Pros & Cons of Alternative Fuels

After examining the potential alternative fuel options, we have chosen engines with performance ratings similar to the Pratt & Whitney PT6A-42 engines. The PT6A-42 has 850shp and weighs 403 lbs. With our TSFC of .5671, our fuel tanks will hold 200 gallons of fuel weighing approximately 1,350 lbs.

8. Constraint Models

In order to estimate the optimum design values for thrust-to-weight ratio and wing loading for the aircraft, constraint diagrams were constructed using limits for important stages of flight. These included takeoff, landing, cruise, ceiling, maximum range, and minimum power. Equations 5.8, 5.9, 5.11, 5.13, and 5.14 (Raymer) were used in these calculations. The design space is found in the area bounded by the constraint curves. Specifically, the thrust-to-weight ratio must be greater than or equal to any point along the takeoff, cruise, and ceiling curves for a given wing loading. Also, the wing loading must be less than or equal to that required for landing, maximum range, and minimum power.

Several preliminary assumptions have also been made in order to construct these diagrams. Because many of the design parameters have yet to be determined, typical values were taken from the course text. These values are given in Table 8 below. Additionally, it was assumed that engine power decreases with altitude in direct proportion to density.

The constraint diagrams for the three initial concepts are given in Figure 22. The constraints for maximum range and minimum power do not appear in the diagram, as the corresponding wing loadings lie far outside that of the landing constraint. The shaded area of the figure is the design range. Any set of thrust-to-weight ratio and wing loading values within this shaded area will satisfy the design requirements for the aircraft.

Turboprop Turbojet

CD0 .02 .015

CLmax 1.5 1.5 e .8 .8

ηp .8 N/A AR 7.6 8 Table 8 - Parameters Used in Constraint Models

Figure 22 - Constraint Diagrams

9. Aircraft Characteristics and Comparisons

Concept 15 Socata TBM 700iv Pilatus PC 12v GTOW 6800 lb 6578 lb 9920 lb W/S (wing loading) 32 lb/ft2 33.9 lb/ft2 lb/ft2 S (wetted wing area) 212 ft2 193.8 ft2 277.8 ft2 AR 7.6 8.97 9.85 b (wing span) 40 ft 1 in 41ft 7 in 52 ft 3 in P/W (power to weight) - props P/W = 0.088 P/W = 0.106 P/W = 0.121

STO 2100 ft 2133 ft 2300 ft

VCruise 250 kts 243 kts 232 kts Acquisition cost $1.65 million $ 2 million $ 2.8 million 100LL - $350/hr $425/hr $400/hr Direct Operating cost Biodiesel - $468/hr Table 9 - Aircraft Comparison

10. Conclusion

A global market exists for a small to medium sized propeller driven aircraft, primarily aimed at operators such as air taxi services, air charter operations, and corporate flight departments. The market suggests a demand in aircraft capable of carrying six passengers and two crew members on a 600 nautical mile trip. A capability of using runways as short as 2,100 ft will also open a multitude of airports providing more convenient point to point service.

The benefit of such an aircraft will be realized not only in the United States, but in the European and Asian markets as well. An initial design mission is drawn from the customer attributes and subsequent analysis with the QFD matrix. Extensive trade studies were used to justify the design requirements and the ratings given in the Pugh’s Method. The aircraft characteristics of the final concept and competing aircraft are shown in Table 9.

Furthermore, the increasing uncertainty about the future of petroleum based fuels will create a need for alternative means of transportation. The new market will need to offer flexible point-to-point air transportation at low cost and high efficiency. In addition, awareness of global climate change will result in a larger demand for more environmentally friendly fuel sources. This study has determined that the aviation industry will be adaptable to such change, and aircraft represented by Concept 15 will meet the needs of a changing market.

11. Appendix:

1. Jet Trade Study Appendix

Figure 1 – Jet Gross Weight vs. Range and Speed ...... 39 Figure 2 – Jet Acquisition Cost vs. Range and Speed ...... 39 Figure 3 – Jet Gross Weight vs. Range and L/Dmax ...... 410 Figure 4 – Jet Acquisition Cost vs. Range and L/Dmax ...... 410 Figure 5 – Jet Gross Weight vs. Range and SFC ...... 421 Figure 6 – Jet Acquisition Cost vs. Range and SFC ...... 421 Figure 7 – Jet Gross Weight vs. Range and Number of Passengers ...... 432 Figure 8 – Jet Acquisition Cost vs. Range and Number of Passengers ...... 432 Figure 9 – Jet Gross Weight vs. Speed and SFC ...... 443 Figure 10 – Jet Acquisition Cost vs. Speed and SFC ...... 443 Figure 11 – Jet Operating Costs/Hour vs. Range and Speed ...... 454 Figure 12 – Jet Operating Costs/Hour vs. Range and SFC ...... 454

J e t - T r a d e S t u d y ( R a n g e a n d S p e e d ) 1 3 5 0 0

1 3 0 0 0 200 kts

1 2 5 0 0

225 kts 1 2 0 0 0 ) s b

l 1 1 5 0 0 250 kts (

t h g i 275 kts e 1 1 0 0 0 W

s 300 kts s 1 0 5 0 0 o r 325 kts G 1 0 0 0 0 350 kts

9 5 0 0

9 0 0 0

8 5 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 6 – Jet Gross Weight vs. Range and Speed

J e t - T r a d e S t u d y ( R a n g e a n d S p e e d ) 4

) 3 M $ (

t s o C 2 . 5 n o i t i s i

u 200 kts q c 2 A 225 kts

250 kts

1 . 5 275 kts 300 kts 325 kts 350 kts 1 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 7 – Jet Acquisition Cost vs. Range and Speed

J e t - T r a d e S t u d y ( R a n g e a n d L / D ) m a x 1 2 5 0 0 10.0 L/D m a x

1 2 0 0 0

1 1 5 0 0

1 1 0 0 0

) 12.5 L/D

s m a x b l (

t 1 0 5 0 0 h g i e

W 15.0 L/D 1 0 0 0 0 m a x s s o r

G 17.5 L/D 9 5 0 0 m a x

9 0 0 0

8 5 0 0

8 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 8 – Jet Gross Weight vs. Range and L/Dmax

J e t - T r a d e S t u d y ( R a n g e a n d L / D ) m a x 3

2 . 4 ) M $ (

2 . 2 t s o C 2 n o i t i s i 1 . 8 u q c A 1 . 6 10.0 L/D m a x 1 . 4

12.5 L/D m a x 1 . 2 15.0 L/D m a x 17.5 L/D 1 m a x 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 9 – Jet Acquisition Cost vs. Range and L/Dmax

J e t - T r a d e S t u d y ( R a n g e a n d S F C ) 1 2 5 0 0 0.8 SFC

1 2 0 0 0

1 1 5 0 0 0.7 SFC

1 1 0 0 0 ) s

b 0.6 SFC l (

t 1 0 5 0 0 h g i e

W 0.5 SFC 1 0 0 0 0 s s o r G 9 5 0 0 0.4 SFC

9 0 0 0 0.3 SFC

8 5 0 0

8 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 10 – Jet Gross Weight vs. Range and SFC

J e t - T r a d e S t u d y ( R a n g e a n d S F C ) 2 . 8

2 . 2 ) M $ (

2 t s o C 1 . 8 n o i t i s i 1 . 6 u

q 0.8 SFC c A 1 . 4 0.7 SFC

0.6 SFC 1 . 2 0.5 SFC 1 0.4 SFC 0.3 SFC

0 . 8 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 11 – Jet Acquisition Cost vs. Range and SFC

J e t - T r a d e S t u d y ( R a n g e a n d N u m b e r o f P a s s e n g e r s )

8 5 0 0 8 pax

8 0 0 0 7 pax

7 5 0 0 6 pax ) s b l

( 5 pax

t 7 0 0 0 h g i e W

s 6 5 0 0 s o r G 6 0 0 0 3 pax

5 5 0 0 2 pax

5 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 12 – Jet Gross Weight vs. Range and Number of Passengers

J e t - T r a d e S t u d y ( R a n g e a n d N u m b e r o f P a s s e n g e r s ) 3 . 5

) 2 . 5 M $ (

t s o C 2 n o i t i s i u q c 1 . 5 A 8 pax 7 pax 6 pax 1 5 pax 4 pax 3 pax 0 . 5 2 pax 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 13 – Jet Acquisition Cost vs. Range and Number of Passengers

J e t - T r a d e S t u d y ( S p e e d a n d S F C )

1 1 5 0 0 200 kts

1 1 0 0 0

225 kts

1 0 5 0 0 250 kts ) s b l (

275 kts

t 1 0 0 0 0 h g i 300 kts e W

325 kts

s 9 5 0 0 s 350 kts o r G 9 0 0 0

8 5 0 0

8 0 0 0 0 . 3 0 . 3 5 0 . 4 0 . 4 5 0 . 5 0 . 5 5 0 . 6 0 . 6 5 0 . 7 0 . 7 5 0 . 8 S F C ( l b / h r / b h p )

Figure 14 – Jet Gross Weight vs. Speed and SFC

J e t - T r a d e S t u d y ( S p e e d a n d S F C ) 4 200 kts

225 kts

) 3 M $

( 250 kts

t s o C 2 . 5 275 kts n o i t i

s 300 kts i u q c 2 325 kts A 350 kts

1 0 . 3 0 . 3 5 0 . 4 0 . 4 5 0 . 5 0 . 5 5 0 . 6 0 . 6 5 0 . 7 0 . 7 5 0 . 8 S F C ( l b / h r / b h p )

Figure 15 – Jet Acquisition Cost vs. Speed and SFC

J e t - T r a d e S t u d y ( R a n g e a n d S p e e d ) 9 0 0

8 5 0 200 kts

7 5 0 225 kts r u

o 7 0 0 H / s

r 250 kts a l l 6 5 0 o D 275 kts 6 0 0

300 kts 5 5 0 325 kts

5 0 0 350 kts

4 5 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 16 – Jet Operating Costs/Hour vs. Range and Speed

J e t - T r a d e S t u d y ( R a n g e a n d S F C ) 7 0 0

0.8 SFC 6 5 0

6 0 0 0.7 SFC r u o H /

s 5 5 0 r 0.6 SFC a l l o D

5 0 0 0.5 SFC

4 5 0 0.4 SFC

0.3 SFC 4 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 17 – Jet Operating Costs/Hour vs. Range and SFC

2. Single Turboprop Trade Study Appendix

Figure 1 – Single Turboprop Gross Weight vs. Range and Speed...... 46 Figure 2 – Single Turboprop Acquisition Cost vs. Range and Speed...... 46 Figure 3 – Single Turboprop Gross Weight vs. Range and L/Dmax...... 47 Figure 4 – Single Turboprop Acquisition Cost vs. Range and L/Dmax...... 47 Figure 5 – Single Turboprop Gross Weight vs. Range and SFC...... 48 Figure 6 – Single Turboprop Acquisition Cost vs. Range and SFC...... 48 Figure 7 – Single Turboprop Gross Weight vs. Range and Number of Passengers...... 49 Figure 8 – Single Turboprop Acquisition Cost vs. Range and Number of Passengers.....49 Figure 9 – Single Turboprop Gross Weight vs. Speed and SFC...... 50 Figure 10 – Single Turboprop Acquisition Cost vs. Speed and SFC...... 50 Figure 11 – Single Turboprop Operating Costs/Hour vs. Range and Speed...... 51 Figure 12 – Single Turboprop Operating Costs/Hour vs. Range and SFC...... 51

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 7 5 0 0 3 5 0 k t s 3 2 5 k t s 3 0 0 k t s 2 7 5 k t s 2 5 0 k t s 7 0 0 0 2 2 5 k t s 2 0 0 k t s ) s b

l 6 5 0 0 (

t h g i e W

s 6 0 0 0 o r G

5 5 0 0

5 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 18 – Single Turboprop Gross Weight vs. Range and Speed

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 2 . 5

3 5 0 k t s

3 2 5 k t s

) 2 3 0 0 k t s M $ (

s 2 7 5 k t s o C

o 2 5 0 k t s i t i s i

u 2 2 5 k t s q c 1 . 5 A 2 0 0 k t s

1 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 19 – Single Turboprop Acquisition Cost vs. Range and Speed

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d L / D ) m a x

9 0 0 0 1 0 . 0 L / D m a x

8 5 0 0

8 0 0 0 )

s 7 5 0 0 b l (

1 2 . 5 L / D t m a x h g i

e 7 0 0 0 W

o 1 5 . 0 L / D r 6 5 0 0 m a x G

1 7 . 5 L / D 6 0 0 0 m a x

5 5 0 0

5 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 20 – Single Turboprop Gross Weight vs. Range and L/Dmax

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d L / D ) m a x 2 . 5

1 0 . 0 L / D m a x

) 2 M $ (

t 1 2 . 5 L / D s m a x o C

n o i t i 1 5 . 0 L / D

s m a x i u q

c 1 7 . 5 L / D 1 . 5 m a x A

1 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 21 – Single Turboprop Acquisition Cost vs. Range and L/Dmax

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 9 0 0 0 0 . 8 S F C 8 5 0 0

8 0 0 0 0 . 7 S F C

7 5 0 0 ) s b l (

0 . 6 S F C t 7 0 0 0 h g i e W 6 5 0 0 0 . 5 S F C s s o r

G 6 0 0 0 0 . 4 S F C

5 5 0 0 0 . 3 S F C

5 0 0 0

4 5 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 22 – Single Turboprop Gross Weight vs. Range and SFC

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 2 . 5

0 . 8 S F C

) 2 0 . 7 S F C M $ (

o 0 . 6 S F C C

n o i t i 0 . 5 S F C s i u q c 1 . 5 0 . 4 S F C A

0 . 3 S F C

1 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 23 – Single Turboprop Acquisition Cost vs. Range and SFC

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d N u m b e r o f P a s s e n g e r s ) 5 5 0 0 8 p a x

5 0 0 0

7 p a x

4 5 0 0

) 6 p a x s b l (

t 4 0 0 0 h g i

e 5 p a x W

s 3 5 0 0 s o r 4 p a x G 3 0 0 0

3 p a x

2 5 0 0 2 p a x

2 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 24 – Single Turboprop Gross Weight vs. Range and Number of Passengers

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d N u m b e r o f P a s s e n g e r s ) 2 . 4

8 p a x 2 . 2

2 7 p a x

1 . 8 ) M

$ 6 p a x (

1 . 6 t s o C 1 . 4 n 5 p a x o i t i s i 1 . 2 u q

c 4 p a x A 1

3 p a x 0 . 8

0 . 6 2 p a x

0 . 4 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 25 – Single Turboprop Acquisition Cost vs. Range and Number of Passengers

T u r b o p r o p - T r a d e S t u d y ( S p e e d a n d S F C ) 8 0 0 0

7 5 0 0 3 5 0 k t s 3 2 5 k t s 3 0 0 k t s 2 7 5 k t s 2 5 0 k t s 2 2 5 k t s 7 0 0 0 2 0 0 k t s ) s b l (

t h g i

e 6 5 0 0 W

s s o r

G 6 0 0 0

5 5 0 0

5 0 0 0 0 . 3 0 . 3 5 0 . 4 0 . 4 5 0 . 5 0 . 5 5 0 . 6 0 . 6 5 0 . 7 0 . 7 5 0 . 8 S F C ( l b / h r / b h p )

Figure 26 – Single Turboprop Gross Weight vs. Speed and SFC

T u r b o p r o p - T r a d e S t u d y ( S p e e d a n d S F C ) 2 . 5 3 5 0 k t s

3 2 5 k t s

3 0 0 k t s

M 2 7 5 k t s $ (

o 2 5 0 k t s C

n o i t

i 2 2 5 k t s s i u q c 1 . 5 2 0 0 k t s A

1 0 . 3 0 . 3 5 0 . 4 0 . 4 5 0 . 5 0 . 5 5 0 . 6 0 . 6 5 0 . 7 0 . 7 5 0 . 8 S F C ( l b / h r / b h p )

Figure 27 – Single Turboprop Acquisition Cost vs. Speed and SFC

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 5 5 0

3 5 0 k t s 5 0 0

3 2 5 k t s 4 5 0

r 3 0 0 k t s u o H /

s 4 0 0 r a l l 2 7 5 k t s o D

3 5 0 2 5 0 k t s

3 0 0 2 2 5 k t s

2 0 0 k t s 2 5 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 28 – Single Turboprop Operating Costs/Hour vs. Range and Speed

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 3 6 0 0 . 8 S F C

0 . 7 S F C

3 4 0 r u

o 0 . 6 S F C H /

s 3 3 0 r a l l o D 0 . 5 S F C 3 2 0

0 . 4 S F C 3 1 0

0 . 3 S F C 3 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 29 – Single Turboprop Operating Costs/Hour vs. Range and SFC

3. Twin Turboprop Trade Study Appendix Figure 1 – Twin Turboprop Gross Weight vs. Range and Speed ...... 54 Figure 2 – Twin Turboprop Acquisition Cost vs. Range and Speed ...... 54 Figure 3 – Twin Turboprop Gross Weight vs. Range and L/Dmax ...... 55 Figure 4 – Twin Turboprop Acquisition Cost vs. Range and L/Dmax ...... 55 Figure 5 – Twin Turboprop Gross Weight vs. Range and SFC ...... 56 Figure 6 – Twin Turboprop Acquisition Cost vs. Range and SFC ...... 56 Figure 7 – Twin Turboprop Gross Weight vs. Range and Number of Passengers ...... 57 Figure 8 – Twin Turboprop Acquisition Cost vs. Range and Number of Passengers ...... 57 Figure 9 – Twin Turboprop Gross Weight vs. Speed and SFC ...... 58 Figure 10 – Twin Turboprop Acquisition Cost vs. Speed and SFC ...... 58 Figure 11 – Twin Turboprop Operating Costs/Hour vs. Range and Speed ...... 59 Figure 12 – Twin Turboprop Operating Costs/Hour vs. Range and SFC ...... 59

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 1 1 5 0 0 3 5 0 k t s 3 2 5 k t s 1 1 0 0 0 3 0 0 k t s 2 7 5 k t s 2 5 0 k t s 2 2 5 k t s 1 0 5 0 0 2 0 0 k t s ) s b l (

t 1 0 0 0 0 h g i e W

s 9 5 0 0 s o r G 9 0 0 0

8 5 0 0

8 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 30 – Twin Turboprop Gross Weight vs. Range and Speed

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 8

6 ) M $ (

s 5 3 5 0 k t s o C

i 3 2 5 k t s t i 4 s i u q

c 3 0 0 k t s A 3 2 7 5 k t s

2 5 0 k t s 2 2 2 5 k t s

2 0 0 k t s 1 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 31 – Twin Turboprop Acquisition Cost vs. Range and Speed

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d L / D ) m a x 1 2 0 0 0

1 0 . 0 L / D m a x 1 1 5 0 0

1 1 0 0 0 )

s 1 0 5 0 0 b l (

h 1 2 . 5 L / D g

i m a x

e 1 0 0 0 0 W

r 9 5 0 0

G 1 5 . 0 L / D m a x

9 0 0 0 1 7 . 5 L / D m a x

8 5 0 0

8 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 32 – Twin Turboprop Gross Weight vs. Range and L/Dmax

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d L / D ) m a x 4 . 5

) 1 0 . 0 L / D 3 . 5 m a x M $ (

t s o C 3 n o i t i s i u q

c 1 2 . 5 L / D 2 . 5 m a x A

1 5 . 0 L / D 2 m a x

1 7 . 5 L / D m a x

1 . 5 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 33 – Twin Turboprop Acquisition Cost vs. Range and L/Dmax

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 1 1 0 0 0 0 . 8 S F C

1 0 5 0 0

0 . 7 S F C 1 0 0 0 0 )

s 9 5 0 0

b 0 . 6 S F C l (

t h g i

e 9 0 0 0 0 . 5 S F C W

r 8 5 0 0

G 0 . 4 S F C

8 0 0 0 0 . 3 S F C

7 5 0 0

7 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 34 – Twin Turboprop Gross Weight vs. Range and SFC

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 4 . 5

) 3 . 5

M 0 . 8 S F C $ (

t s o C 3 n o i

t 0 . 7 S F C i s i u q c 2 . 5 A 0 . 6 S F C

0 . 5 S F C 2

0 . 4 S F C

1 . 5 0 . 3 S F C 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 35 – Twin Turboprop Acquisition Cost vs. Range and SFC

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d N u m b e r o f P a s s e n g e r s ) 9 0 0 0

8 p a x 8 5 0 0

8 0 0 0 7 p a x

7 5 0 0 6 p a x ) s b

l 7 0 0 0 (

h 5 p a x g i

e 6 5 0 0 W

s 4 p a x s 6 0 0 0 o r G

5 5 0 0 3 p a x

5 0 0 0 2 p a x 4 5 0 0

4 0 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 36 – Twin Turboprop Gross Weight vs. Range and Number of Passengers

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d N u m b e r o f P a s s e n g e r s ) 5

) 3 . 5 M $ (

C 8 p a x

n o i t i 2 . 5 7 p a x s i u q c 2 6 p a x A

5 p a x 1 . 5 4 p a x

1 3 p a x 2 p a x 0 . 5 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 37 – Twin Turboprop Acquisition Cost vs. Range and Number of Passengers

T u r b o p r o p - T r a d e S t u d y ( S p e e d a n d S F C ) 1 3 0 0 0

3 5 0 k t s 1 2 5 0 0 3 2 5 k t s 3 0 0 k t s 2 7 5 k t s 1 2 0 0 0 2 5 0 k t s 2 2 5 k t s 2 0 0 k t s 1 1 5 0 0 ) s b

l 1 1 0 0 0 (

t h g i

e 1 0 5 0 0 W

s 1 0 0 0 0 o r G 9 5 0 0

9 0 0 0

8 5 0 0

8 0 0 0 0 . 3 0 . 3 5 0 . 4 0 . 4 5 0 . 5 0 . 5 5 0 . 6 0 . 6 5 0 . 7 0 . 7 5 0 . 8 S F C ( l b / h r / b h p )

Figure 38 – Twin Turboprop Gross Weight vs. Speed and SFC

T u r b o p r o p - T r a d e S t u d y ( S p e e d a n d S F C ) 9

3 5 0 k t s 8

7 3 2 5 k t s ) M

t 3 0 0 k t s s o C 5 n o

i 2 7 5 k t s t i s i u

c 2 5 0 k t s A

3 2 2 5 k t s

2 0 0 k t s 2

1 0 . 3 0 . 3 5 0 . 4 0 . 4 5 0 . 5 0 . 5 5 0 . 6 0 . 6 5 0 . 7 0 . 7 5 0 . 8 S F C ( l b / h r / b h p )

Figure 39 – Twin Turboprop Acquisition Cost vs. Speed and SFC

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S p e e d ) 6 4 0 2 0 0 k t s

6 2 0 2 2 5 k t s

2 5 0 k t s

2 7 5 k t s 6 0 0 3 0 0 k t s

r 3 2 5 k t s

u 3 5 0 k t s o H /

s 5 8 0 r a l l o D

5 2 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 40 – Twin Turboprop Operating Costs/Hour vs. Range and Speed

T u r b o p r o p - T r a d e S t u d y ( R a n g e a n d S F C ) 1 0 0 0 0 . 8 S F C

0 . 7 S F C 8 0 0 r u o H

/ 0 . 6 S F C

s 7 0 0 r a l l o D 0 . 5 S F C 6 0 0

0 . 4 S F C 5 0 0 0 . 3 S F C

4 0 0 4 0 0 4 5 0 5 0 0 5 5 0 6 0 0 6 5 0 7 0 0 7 5 0 8 0 0 R a n g e ( n m )

Figure 41 – Twin Turboprop Operating Costs/Hour vs. Range and SFC

4. MATLAB and FLOPS Code Appendix

61. Jet Parameter Setup Function 621. Single Turbo Parameter Setup Function 62. Design Mission Weight Fraction Analyses 64. Constraint Diagram Functions 66. Trade Study Plot Functions 66. FLOPS Input Data – Twin Turboprop 67. FLOPS Input Data – Single Turboprop 68. FLOPS Input Data – VLJ

Note: The functions that generate the plots are long and repetitive, so they are not included in this appendix. They are available in the AAE 451 restricted access folder for Group 6 on the course webpage, in the folder labelled SDR MATLAB Code.

1. Jet Parameter Setup

% Jet Aircraft clear all clc

% Aircraft Parameters Type = 1; % (0 = Prop) (1 = Jet) np = 6; % Number of Passengers nc = 2; % Number of Crew L_Dm = 15; % L/Dmax (Raymer) Cc = 0.7; % Cruise SFC (Raymer) Cl = 0.6; % Loiter SFC (Raymer) w0_g = 9000; % Gross Weight Estimate (lbs) ar = 8.0; % Aspect Ratio e = 0.8; % Oswald Efficiency cdo = 0.015; % Zero Lift Drag c = Cc; % SFC clmax = 1.5; % CLmax altc = 30000; % Cruise Altitude u = 0.3; % Landing Rolling Friction

% Empty Weight Fraction we/w0 = A*w0^B A = 1.59; % Raymer B = -0.10; % Raymer

% Cost Model Constants [Cost = (w0^a)*(R^b)*(V^c)*exp(d)] a = 1.4532; b = -1.4285; c = -0.4833; d = 12.9303;

% dollars per flight hour a2 = 1.3726; b2 = -0.3169; c2 = -0.3331; d2 = -2.3700;

Reg(1) = A; Reg2(1) = A; Reg(2) = B; Reg2(2) = B; Reg(3) = a; Reg2(3) = a2; Reg(4) = b; Reg2(4) = b2; Reg(5) = c; Reg2(5) = c2; Reg(6) = d; Reg2(6) = d2;

% Mission Parameters V = 350; % Speed (knots) R = 600; % Range(nm) E = 45; % Loiter Time(min)

% Crew and Passenger Weights wc = 150; % Crew Weight (per person) (lbs) wct = wc*nc; wp = 200; % Passenger Weight (per person) (lbs) wpt = wp*np; wt = wct + wpt; % Total Crew and Passenger Weight (lbs)

[w0,wf,co] = mission(R,Cc,E,Cl,L_Dm,w0_g,V,wt,Type,Reg); fprintf('\nAircraft Gross Weight: %.0f lbs\n', w0) fprintf('\nAircraft Acquisistion Cost: $%.3f million\n\n', co);

2. Single Prop Parameter Setup

% Single Prop Aircraft clear all

% Aircraft Parameters Type = 0; % (0 = Prop) (1 = Jet) np = 6; % Number of Passengers nc = 2; % Number of Crew L_Dm = 15; % L/Dmax (Raymer) Cc = 0.7; % Cruise SFC (Raymer) (lb/hr/bhp) Cl = 0.6; % Loiter SFC (Raymer) (lb/hr/bhp) w0_g = 9000; % Gross Weight Estimate (lbs) ar = 7.6; % Aspect Ratio e = 0.8; % Oswald Efficiency cdo = 0.02; % Zero Lift Drag c = Cc; % SFC clmax = 1.5; % CLmax altc = 20000; % Cruise Altitude u = 0.3; % Landing Rolling Friction

% Empty Weight Fraction [we/w0 = A*w0^B] A = 0.2705; % Regression B = 0.0792; % Regression

% Cost Model Constants [Cost = (w0^a)*(R^b)*(V^c)*exp(d)] a = 1.0693; b = 0.1149; d = 0.2657; c = 0.7100;

% dollars per flight hour a2 = 0.3762; b2 = -0.0899; c2 = 1.2038; d2 = -3.5266;

Reg(1) = A; Reg2(1) = A; Reg(2) = B; Reg2(2) = B; Reg(3) = a; Reg2(3) = a2; Reg(4) = b; Reg2(4) = b2; Reg(5) = c; Reg2(5) = c2; Reg(6) = d; Reg2(6) = d2;

% Mission Parameters V = 240; % Speed (knots) R = 600; % Range(nm) E = 45; % Loiter Time(min)

% Crew and Passenger Weights wc = 150; % Crew Weight (per person) (lbs) wct = wc*nc; wp = 200; % Passenger Weight (per person) (lbs) wpt = wp*np; wt = wct + wpt; % Total Crew and Passenger Weight (lbs)

[w0,wf,co] = mission(R,Cc,E,Cl,L_Dm,w0_g,V,wt,Type,Reg); fprintf('\nAircraft Gross Weight: %.0f lbs\n', w0) fprintf('\nAircraft Acquisistion Cost: $%.3f million\n\n', co);

3. Design Mission Weight Fraction Analysis function [w0,wf,co] = mission(R,Cc,E,Cl,L_Dm,w0_g,V,wt,Type,Reg);

% Unit Conversions V = V.*1.6878; % Speed (ft/sec) R = R.*6076.1155; % Range (ft) E = E.*60; % Loiter Time(sec)

% Aircraft Type if (Type == 0) % Propeller Aircraft np = 0.8; % Propeller Efficiency (Raymer) L_Dc = L_Dm; % Cruise L/D (Raymer) L_Dl = 0.866.*L_Dm; % Loiter L/D (Raymer) Cc = Cc.*V./(550.*np)./3600; % Cruise SFC (1/sec) Cl = Cl.*V./(550.*np)./3600; % Cruise SFC (1/sec) M = 0.325; % Cruise Mach Number else % Turbofan Aircraft L_Dc = 0.866.*L_Dm; % Cruise L/D (Raymer) L_Dl = L_Dm; % Loiter L/D (Raymer) Cc = Cc./3600; % Cruise SFC (1/sec) Cl = Cl./3600; % Cruise SFC (1/sec) M = 0.595; % Cruise Mach Number end w0_1 = 0; w0_2 = w0_g; while (abs(w0_2-w0_1)> 1.0) w0_1 = w0_2; w1_w0 = 0.970; % Warmup and Takeoff (Raymer) w2_w1 = 1.0065-0.0325*M; % Climb (Raymer) 6.9 w3_w2 = exp(-R.*Cc./V./L_Dc); % Berguet Range Equation w4_w3 = 0.990; % Descent (Raymer) w5_w4 = 0.990; % Missed Approach (Raymer) w6_w5 = 0.985; % Climb (Raymer) w7_w6 = 0.960; % 200nm Cruise (Raymer) w8_w7 = exp(-E.*Cl./L_Dl); % Endurance Equation w9_w8 = 0.990; % Descent (Raymer) w10_w9 = 0.995; % Landing (Raymer) w10_w0 = w1_w0.*w2_w1.*w3_w2.*w4_w3.*w5_w4.*... w6_w5.*w7_w6.*w8_w7.*w9_w8.*w10_w9; wf_w0 = 1.01.*(1-w10_w0); % Fuel Fraction we_w0 = Reg(1).*w0_1.^Reg(2); % Gross Weight (Estimated) w0_2 = (wt)./(1-wf_w0-we_w0); % Gross Weight (Calculated) end w0 = w0_2; % Gross Weight (Converged) wf = wf_w0.*w0; % Fuel Weight V = V./1.6878; R = R./6076.1155; co = (w0.^Reg(3)).*(R.^Reg(4)).*(V.^Reg(5)).*exp(Reg(6)); co = co./(1e6); % Cost ($Millions) end

4. Constraint Analysis Generation

% Constraint Analysis g=32.2; rhosl=.002378;pi=3.14159; wos= linspace(1,150,150); k = 1/(pi*ar*e); emax = 1/(2*sqrt(k*cdo));

% Take-Off Constraint beta = 1; alt = 0; d = 2100; sigma = density(alt); alpha = thrust(alt); towto = (1.1^2)*(beta^2.)*wos/(alpha*sigma*rhosl*g*clmax*d);

% Cruise Speed Constraint beta = 1; alt = altc; v = V; v = v*1.6878; sigma = density(alt); alpha = thrust(alt); q = 0.5*sigma*rhosl*(v^2.); towc = (beta/alpha)*(q*cdo./(beta*wos)+ k*beta*wos/q);

% Cruise Altitude Constraint beta = 1; alt = altc; sigma = density(alt); alpha = thrust(alt); towa = beta/(emax*alpha)*ones(size(wos));

% Landing beta = 1; alt = 0; sigma = density(alt); rhol = rhosl*sigma; wosl = g*d*rhol*clmax*u/((1.15^2)*beta); towl = linspace(0,0.5,500); figure() hold all plot(wos,towto) plot(wos,towc) plot(wos,towa) plot(wosl,towl,'-c') axis([0 75 0 0.5]) title(' Constraint Analysis ') xlabel(' W/S - Wing Loading (lb/ft^2)') ylabel(' T/W - Thrust to Weight Ratio ') legend('Take-Off','Cruise Speed','Cruise Altitude','Landing') grid on; function alpha=thrust(h) % **** Thrust model **** % returns the ratio of thrust to sea level thrust alpha = density(h);

function sigma=density(altitude) % note sigma returns sigma ratio - sigma % calculate properties for the standard atmosphere Ts=518.69; rhos=.0023769; Ps=2116.2; g0=32.2; R=1716.; a1=-.00356; h1=36150; T1=389.99; a2=.001631; h2=82300; if (altitudeh1 & altitudeh2) T= T1 + a2*(altitude-h2); sigma1=(T1/Ts).^(-(g0/(a1*R)+1)); sigma2=exp(-g0/(R*T1)*(h2-h1))*sigma1; sigma=(T/T1).^(-(g0/(a2*R)+1))*sigma2; return end end

5. Trade Study Plot Generation

% Trade Studies close all

% Plot Generation (0 = No) (1 = Yes) Plot1 = 0; % Trade Study (Range and Speed) Plot2 = 0; % Trade Study (Range and L/Dmax) Plot3 = 0; % Trade Study (Range and SFC) Plot4 = 0; % Trade Study (Range and Number of Passengers) Plot5 = 0; % Trade Study (Speed and SFC) Plot6 = 0; % Trade Study (Fuel Weight vs. Range and Speed) Plot7 = 0; % Trade Study (DOC vs. Range and Speed) Plot8 = 0; % Trade Study (DOC vs. Range and SFC) Plot20 = 0; % Variation Study

% Variation of Parameters Rs = linspace(400,800,20); % Range(nm) Vs = 200:25:350; % Speed (knots) Ccs = 0.3:0.1:0.8; % Cruise SFC (lb/hr/bhp) L_Dms = 10:2.5:17.5; % L/Dmax nps = 2:1:8; % Number of Passengers wp = 200; % Passenger Weight (per person) (lbs) wpts = nps.*wp; wts = wct + wpts; % Total Crew and Passenger Weight (lbs)

% Trade Studies if(Plot1 == 1);rv(Rs,Cc,E,Cl,L_Dm,w0_g,Vs,wt,Type,Reg);else;end; if(Plot2 == 1);rld(Rs,Cc,E,Cl,L_Dms,w0_g,V,wt,Type,Reg);else;end; if(Plot3 == 1);rSFC(Rs,Ccs,E,Cl,L_Dm,w0_g,V,wt,Type,Reg);else;end; if(Plot4 == 1);rwt(Rs,Cc,E,Cl,L_Dm,w0_g,V,wts,Type,Reg);else;end; if(Plot5 == 1);vSFC(R,Ccs,E,Cl,L_Dm,w0_g,Vs,wt,Type,Reg);else;end; if(Plot6 == 1);w_wf(Rs,Cc,E,Cl,L_Dm,w0_g,Vs,wt,Type,Reg);else;end; if(Plot7 == 1);doc_rv(Rs,Cc,E,Cl,L_Dm,w0_g,Vs,wt,Type,Reg2);else;end; if(Plot8 == 1);doc_rSFC(Rs,Ccs,E,Cl,L_Dm,w0_g,V,wt,Type,Reg2);else;end; % Variation Study N = 100; % Number of Variations per Parameter dv = 0.5; % Variation Range (+/-) Fraction dvh = 1+dv; dvl = 1-dv; Rs = linspace(dvl*R,dvh*R,N); % Range(nm) Vs = linspace(dvl*V,dvh*V,N); % Speed (knots) Ccs = linspace(dvl*Cc,dvh*Cc,N); % Cruise SFC (lb/hr/bhp) L_Dms = linspace(dvl*L_Dm,dvh*L_Dm,N); % L/Dmax nps = linspace(dvl*np,dvh*np,N); % Number of Passengers wpts = nps.*wp; wts = wct + wpts; % Total Crew and Passenger Weight (lbs) wt = [wt wct]; if(Plot20 == 1); var(R,Rs,V,Vs,Cc,Ccs,L_Dm,L_Dms,wt,wts,E,Cl,w0_g,Type,Reg); else;end;

% Note: The functions that generate these plots are long and repetitive, so they are not included in this appendix. They are available in the AAE 451 restricted access folder for Group 6 on the course webpage, in the folder labeled SDR Matlab Code.

FLOPS Model Generation: Twin Turboprop Input

TWIN TURBOPROP EXAMPLE for AAE 451, Spring 2006 Run a full analysis including costs - caution, costs may not be accurate for twin turboprop $OPTION IOPT=1, IANAL=3, ICOST=1, $END Enter fuselage dimensions assume all pax are first class, use two wing-mounted engines Tails are specified with volume coefficients and default parameters Empty weight reduced, appears FLOPS equations over-predict twin turboprop $WTIN WF=5.0, DF=5.0, XL=35.5, FPITCH=24., NFABR=1, NTABR=0, NEW=2, NEF=0, FULFMX=2573.,NPF=6, NPT=0, NFLCR=2, NSTU=0, NGALC=0, THRSO=550., EWMARG=-0.1,FCOMP=1., $END Maintain constant wing loading, thrust/weight ratio, and modified tail volume coefficients based on existing twin turboprop. $CONFIN GW=10000.0, DESRNG=800., AR=8.5, WSR=35., TWR=.30, TCA=.14, TR=0.44, SWEEP=0., HTVC=0.975, VTVC=0.1, VCMN=0.5, CH=5000., $END Moderate technology wing $AERIN AITEK=1., FLTO=2100., $END Calculate cost information, starting development year 2006, fuel price Feb 2006 use 100 percent first class seating, production run 300 a/c $COSTIN DEVST=2006., DYEAR=2006, FUELPR=3.50, NPOD=2, PLMQT=2014., Q=300., $END Generate engine deck in cycle analysis module and extrapolate to get consistent flight idle data $ENGDIN IDLE=1, IGENEN=1, MAXCR=1, NGPRT=0, $END Generate a turboprop engine, use default prop performance via ETAPRP $ENGINE IENG=4, IPRINT=0, OPRDES=29.5, TETDES=2660.0, ETAPRP=0.840, SHPOWA=60.0, $END Size aircraft for specified range, fly minimum fuel-to-climb, optimum altitude for cruise Mach, and max L/D descent $MISSIN IFLAG=2, IRW=1, TAXOTM=10., TAKOTM=0.4, TAXITM=10., TIMMAP=5., ITTFF=1, FWF=-1., RESRFU=0.05, THOLD=.05, $END START CLIMB CRUISE DESCENT

Single Turboprop Input

SINGLE TURBOPROP EXAMPLE for AAE 451, Spring 2006 Run a full analysis including costs - caution, costs may not be accurate for twin turboprop $OPTION IOPT=1, IANAL=3, ICOST=1, $END Enter fuselage dimensions assume all pax are first class, use two wing-mounted engines Tails are specified with volume coefficients and default parameters Empty weight reduced, appears FLOPS equations over-predict twin turboprop $WTIN WF=5.0, DF=5.0, XL=47.25, FPITCH=24., NFABR=1, NTABR=0, NEW=0, NEF=1, FULWMX=23.,NPF=6, NPT=0, NFLCR=2, NSTU=0, NGALC=0, THRSO=500., EWMARG=-.05,FCOMP=1.,NETAW = 1,HHT = 1., $END Maintain constant wing loading, thrust/weight ratio, and modified tail volume coefficients based on existing twin turboprop. $CONFIN GW=10000.0, DESRNG=600., AR=10.2, WSR=35., TWR=.195, TCA=.14, TR=0.44, SWEEP=0., HTVC=0.975, VTVC=0.1, VCMN=0.4, CH=20000., $END Moderate technology wing $AERIN AITEK=1.5, FLTO=2100., $END Calculate cost information, starting development year 2006, fuel price Feb 2006 use 100 percent first class seating, production run 300 a/c $COSTIN DEVST=2006., DYEAR=2006, FUELPR=3.50, NPOD=2, PLMQT=2014., Q=300., $END Generate engine deck in cycle analysis module and extrapolate to get consistent flight idle data $ENGDIN IDLE=1, IGENEN=1, MAXCR=1, NGPRT=0, $END Generate a turboprop engine, use default prop performance via ETAPRP $ENGINE IENG=4, IPRINT=0, OPRDES=29.5, TETDES=2660.0, ETAPRP=0.840, SHPOWA=60.0, $END Size aircraft for specified range, fly minimum fuel-to-climb, optimum altitude for cruise Mach, and max L/D descent $MISSIN IFLAG=2, IRW=1, TAXOTM=10., TAKOTM=0.4, TAXITM=10., TIMMAP=5., ITTFF=1, FWF=-1., RESRFU=0.05, THOLD=.05, $END START CLIMB CRUISE DESCENT END

Very Light Jets Input

SUBSONIC VL JET EXAMPLE for AAE 451, Spring 2006 Run a full analysis including costs $OPTION IOPT=1, IANAL=3, ICOST=1, $END Enter fuselage dimensions assume all pax are first class, use two fuselage-mounted engines Tails are specified with volume coefficients and default parameters $WTIN WF=6.0, DF=6.0, XL=50.0, FPITCH=24., NFABR=1, NTABR=0, NEF=2, FULWMX=29.,NPF=6, NPT=0, NFLCR=2, NSTU=0, NGALC=0, THRSO=7000., $END Maintain constant wing loading, thrust/weight ratio, and modified tail volume coefficients. $CONFIN GW=18000.0, DESRNG=2000., AR=8., WSR=50., TWR=.40, TCA=.11, TR=0.27, SWEEP=25., HTVC=1.6, VTVC=0.2, VCMN=0.74, CH=41000., $END Moderate technology wing $AERIN AITEK=1.5, FLTO=2100., $END Calculate cost information, starting development year 2006, fuel price Feb 2006 use 100 percent first class seating, production run 300 a/c $COSTIN DEVST=2006., DYEAR=2006, FUELPR=3.50, NPOD=2, PLMQT=2014., Q=300., $END Generate engine deck in cycle analysis module and extrapolate to get consistent flight idle data $ENGDIN IDLE=1, IGENEN=1, MAXCR=1, NGPRT=0, $END Generate a separate flow turbofan with two compressor components and let the optimum bypass ratio be computed $ENGINE IENG=2, IPRINT=0, OPRDES=29.5, FPRDES=1.67, TETDES=2660.0, $END Size aircraft for specified range, fly minimum fuel-to-climb, optimum altitude for cruise Mach, and max L/D descent $MISSIN IFLAG=2, IRW=1, TAXOTM=10., TAKOTM=0.4, TAXITM=10., TIMMAP=5., ITTFF=1, FWF=-1., RESRFU=0.05, THOLD=.05, $END START CLIMB CRUISE DESCENT END

71 i12. References

Eclipse 500 specifications: http://www.eclipseaviation.com/eclipse_500/specifications/ ii Citation Mustang specifications: http://mustang.cessna.com/site.chtml iii Pilatus PC12 specifications: http://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3 iv Corporate aursearch international inc http://www.caijets.com/tbm_performance_takeoff.php v Wikipedia http://www.answers.com/topic/pilatus-pc-12

13. Merit Pool Team MemberMeritJohn Collins15.5Chad Davis13.5Chris Fles12.5Danny Sze Ling Lim14.5Justin Rohde13.5Ryan Schulz0Ronald Wong15.5Yusaku Yamashita15.0Total100