AAE 451 UAV PROPOSAL SYSTEMS DEFINITION REVIEW DOCUMENTATION TEAM 4

Transcription

1 AAE 451 UAV PROPOSAL SYSTEMS DEFINITION REVIEW DOCUMENTATION TEAM 4 Kevin Kwan Dan Pothala Mohammed Abdul Rahim Niole Risley Sara Tassan John Thornton Sean Wook Alvin Yip April 5, 2007 Page 1 of 72

2 Table of Contents Table of Contents. 2 Exeutive Summary Introdution Produt Definition Business Case Reap Strategy Competition Market Size Market Outlook Cost Analysis Conept Seletion Initial Conepts Pugh s Method Seleted Conept External Layout Internal Layout Payload Integration Constraint Analysis Constraints Take off Constraints Sustained Turn Constraint Constraint Results Sizing Fuel Fration for Cruise and Loiter Carpet Plot Aspet Ratio Analysis Aerodynami Analysis Airfoil Seletion Wing Sweep Taper Twist Propulsion System Engine Seletion Propeller Sizing Stability Analysis Stability Definition Neutral Point Calulation Summary. 65 R.0 Referenes.. 66 A.0 Appendix A 69 April 5, 2007 Page 2 of 72

3 EXECUTIVE SUMMARY: Throughout the world today, there is an inreased demand for Unmanned Aerial System (UAS) in many different industries for many different purposes. A partiularly high need has developed for UAVs with ontinuous area overage apabilities. There are several industries in whih these vehiles would be useful; however, few options are available for the ustomer at an affordable prie. In reent years, there has been a realization among airraft manufaturers and the publi in general of the huge potential that exists in a ivilian UAS market. Law enforement and news agenies, with their heliopter fleets, have to deal with aquisition osts in the millions, and operating osts in the thousands of dollars every hour. The market is ready for the introdution of an Unmanned Aerial System that an provide most of the advantages of a heliopter, while providing huge ost savings and eliminating the risks of putting a rew in the air. The Metro-Sout UAS is being designed speifially around payload pakages suh as the Cineflex V14 high resolution aerial TV amera, and the ThermaCam SC3000, both of whih are reognized as top of the line in their respetive ategories of news overage and surveillane. In addition, ustomers an pik a amera of hoie to use as long as it meets weight, size, and mount adaptability requirements for the Metro-Sout. The Metro-Sout will be apable of arrying two similar, but alternate payload load-outs based on the mission requirements. For instane, a ustomer using the Metro-Sout for news overage would be able to arry single high-resolution TV amera weighing 67 lbs, while a ustomer using the Metro-Sout for news overage would be able to arry a payload pakage omprised of a low to moderate resolution surveillane day/night/ir amera, a still-shot amera, and a radar gun weighing 64 lbs. The Metro-Sout will be able to take-off from small airports in runway distanes as short as 1000 feet. It an travel as far as a 150 miles, and loiter on-station for 5 hours. Current estimates put the gross take-off weight at 600 lbs. In addition, the Metro-Sout UAS will have ompetitive aquisition and operation osts, far below the urrent average for news and law enforement heliopters, and on par with similar UAVs. Team 4 s goal with the Metro-Sout UAS is to provide news overage and law enforement ustomers with a intelligently and ollaboratively-designed ompetitive, effiient UAS to ahieve mission objetives with less risk, less expenditure, and a greater degree of ustomer satisfation. April 5, 2007 Page 3 of 72

4 1.0 Introdution 1.1 Produt Definition The Metro-Sout will be a remotely flown, multi-purpose Unmanned Aerial System designed to operate over highly populated metropolitan environments safely and quietly, in support of the ativities of various news agenies and law enforement departments nationwide. It will be designed to perform ontinuous area overage and meet all urrent FAA regulations for an Unmanned Aerial System operating in airspae over urban areas in the United States. The aerial ativities of news and law enforement agenies have mission requirements that have distintly different goals, but similar harateristis in terms of endurane, range, and many other things. Eah type of ustomer would require different payload load-outs for their type of mission. However, both types of mission load-outs have similar payload weights and set-ups, and therefore, a single airframe an satisfy the mission requirements of both types of ustomers. 2.0 Business Case 2.1 Strategy Team 4 s business strategy is targeted primarily at providing a ost-effetive UAS alternative to manned heliopters for many news agenies and law enforement departments. These agenies have traditionally relied on heliopters for their operations, and inur large expenses in aquisition and operating osts. To suessfully ompete, the UAS would need to offer more than just a ompetitive ability to onvine the ustomer to aquire a new, unproven aerial vehile and replae or omplement the urrent and proven one. 2.2 Competition Aording to market studies onduted by an industry wathdog, (Heliopter International Assoiation), the five year average for for new turbine heliopter sales to law enforement agenies has been 40.2 per year making up an average market share of 26.7 % for new turbine heliopter sales. Figures 2.1, 2.2 and 2.3 below hart the trends in the April 5, 2007 Page 4 of 72

5 numbers of heliopter sales to law enforement and US publi servie agenies over the previous ten years. New Turbine Heliopter Sales to Law Enforement (Last 10 years) Sales Market Share (%) Year Figure 2.1: 10 year trend for new turbine heliopter sales to law enforement agenies in the US [1.4] Figure 2.2: Annual Turbine Heliopter Sales growth to US publi servie agenies April 5, 2007 Page 5 of 72

6 Figure 2.3: Estimated urrent fleet of piston Heliopters in US law enforement [1.3] The two most popular types of rotorraft in use by news agenies are the Bell 206 Jetranger and the Euroopter AS-350. Other rotorraft gaining in popularity are the MD H-500, and Robinson R-44 due to their relative ost-effetiveness. [1.5] Table 2.1, and figures 2.4 and 2.5 below show the trends for the aquisition and operating osts for various popular ompeting heliopters in urrent use. Airraft Aquisition Costs ($) Hourly Operating Costs ($) Bell 206B Jetranger 1,200, Euroopter AS-350 1,670, MD H-500 (used 1981) 475, Robinson R , Soure: Table 2. 1: Competing Heliopter Aquisition and Operating Costs April 5, 2007 Page 6 of 72

7 Heliopter Aquisition Costs US Dollars $1,800,000 $1,600,000 $1,400,000 $1,200,000 $1,000,000 $800,000 $600,000 $400,000 $200,000 $0 Bell 206B Jetranger Euroopter AS-350 MD H-500 (used 1981) Robinson R-44 Aquistion Costs Figure 2.4: Aquisition Costs for various ompeting heliopters urrently in use [6] Competing Heliopters' Hourly Operating Costs $900 $800 $700 US dollars $600 $500 $400 $300 $200 $100 Hourly Operating Costs $0 Bell 206B Jetranger Euroopter AS-350 MD H-500 (used 1981) Robinson R- 44 Figure 2.5: Hourly Operating Costs for various ompeting heliopters urrently in use [6] The Metro-Sout UAS s biggest advantage will be ost. By replaing the heliopter with the Metro-Sout, or supplementing urrent heliopter operations to lower heliopter flight hours, ustomers will save in operating and maintenane osts. April 5, 2007 Page 7 of 72

8 Team 4 has set an initial aquisition ost of $350,000 for the Metro-Sout UAS, as this is most ompetitive prie manageable to maintain a deent profit margin, and produt quality to ompete with heliopters urrently used for similar missions. Additionally, Team 4 has established target yearly operating osts for the Metro-Sout UAS of around $10,000 a year. 2.3 Market Size To estimate the market size, the team researhed the existing number of heliopters operating in the United States. Aording to the Airraft Owners and Pilots Assoiation (AOPA) of Ameria, there are approximately 7,000 ivilian heliopters operating in the US. Of these, about 5,000 are privately owned. This was onfirmed by the rotor.om marketplae reports. The rotor.om market letters reported that law enforement agenies in the US had a total of 154 onfirmed ative piston heliopters, and 483 onfirmed ative turbine heliopters in use in the Fall of These numbers are representative of new heliopters sold to publi servie agenies in the past 20 years and do not take into aount the 797 military surplus turbine heliopters that US publi servie/law enforement agenies aquired from the US army between 1993 and However, the rotor.om reports found that most of these military surplus heliopters have been serving as parts supply for operational heliopters, and Team 4 annot derive a reasonable estimate for the number of ative military surplus heliopters used in law enforement. The net total for heliopters used by law enforement today is in the range. [1.3], [1.4] Eletroni news eathering is the other target market for the team s produt. The two most popular types of rotorraft in use by news agenies are the Bell 206 Jetranger and the Euroopter AS-350. Other rotorraft gaining in popularity are the MDonnell Douglas H- 500, and Robinson R-44 due to their relative ost-effetiveness. Team 4 estimated that the total number of ative heliopters involved in news gathering was somewhere in the range. The team arrived at this estimate by looking at the number of major ities that might require news gathering heliopters. The team estimated onservatively about 100 ities in the US that would have a news market large enough for news heliopters. Also, the team estimated additionally that there are about 40 metropolitan areas in the US large enough to have two or more news agenies with a heliopter on stand-by at all times. [1.5] April 5, 2007 Page 8 of 72

9 From this, the team estimated that law enforement and news agenies ombined were flying nearly 2000 heliopters in the US. This in itself is a sizeable market. However, the relative ease of operability and ost-effetiveness of the Metro-Sout design should make it attrative to a number of smaller News Agenies and law enforement agenies that annot afford the ost of owning and operating heliopters. [1.2] 2.4 Market Outlook Team 4 attempted to develop a realisti predition of the sales trend for the Metro-Sout Unmanned Aerial System over the ourse of its expeted design life. The team estimated that the Metro-Sout UAS would need a minimum of 3 years for design, testing, and ertifiation. Prodution is slated to start in the year 2010, with an initial sales estimate of 20 airframes. Team 4 estimated the sales trend from 2010 through 2021 on a lassi model desribing the introdution of new tehnology into the industry. Sales are estimated to inrease nearly exponentially after the first year, as ustomers reognize the potential in ost-savings of aquiring the Metro-Sout UAS. The expeted sales growth an be attributed to a number of news agenies and law enforement agenies phasing out portions of their aging heliopter fleets in favor of the UAS. Also, the team expets smaller markets to open up in smaller metropolises in about 5-6 years. At this point, the market for the Metro-Sout model should have realized its full potential. Full-sale prodution of about a 150 units will only our around the year 2015, and sales will regress soon after. The reason for the regression is that other manufaturers will ath on to the market potential, and the Metro-Sout will start to beome a less ompetitive option. The team expets this to our in the period. However the suess of the first model would allow Team 4 to develop a newer more ompetitive UAS to ompete with UASs design by other airraft manufaturers as they ome out. Table 2.2 below lists the expeted sales figures in the period for the Metro-Sout UAS. Figure 2.6 depits the trend of marketplae tehnology aeptane, sales growth, realization of full potential, and sales regression expeted over the next deade for the Metro-Sout produt. April 5, 2007 Page 9 of 72

10 Market Outlook (Expeted Sales Figures) Year Metro-Sout Expeted Sales Expeted Market Share (%) Competing Heliopters Sold Competing UASs sold Total Market Size Total Sales Yearly Average Table 2. 2: Antiipated Market Outlook for the Metro-Sout UAV Market Outlook trend for the Metro-Sout Year Figure 2.6: Expeted market trends for Metro-Sout Sales Expeted Market Share (%) Expeted Sales Figures April 5, 2007 Page 10 of 72

11 As seen from figure 2.6, produt aeptane in the marketplae is expeted to our in the period From thereon, sales growth is expeted to our linearly till the full sales potential is reahed around Sales are expeted to then regress as the market reeives an influx of newer more ompetitive UAVs. However, sales are still expeted to ontinue to smaller news agenies and law enforement agenies due to the Metro-Sout s relative inexpensiveness by the period Figure 2.7 below highlights Team 4 s expeted general marketplae trends for the growth of the UAS market among our target ustomers. It is expeted that the size of the UAS market will grow linearly before stabilizing around the year Market Outlook trend for the target market Expeted Market Share (%) Expeted Sales Figures Competing UASs Sold Year Total Market Size Figure 2.7: Expeted Market Trends for UAS sales to Law Enforement and News Agenies 2.5 Cost Analysis Team 4 reiterates that these are expeted sales figures, and should be onsidered only potentially representative of the final suess of the produt. The final step of the team business ase was to estimate total monetary returns on the Metro-Sout projet. The team formulated an expression for produt value depreiation over the years, and inorporated this into the monetary outlook. The team antiipates a net profit of around $105 million at April 5, 2007 Page 11 of 72

12 the end of 2021, from the Metro-Sout program. The details of the monetary return and profits generated on all sales are listed below in Table 2.3. Year Metro-Sout Expeted Sales Expeted Prie Tag ($) with depreiation Monetary Return ($) Unit Prodution Cost ($) Prodution Costs ($) , ,000, ,670, , ,500, ,175, , ,500, ,762, , ,000, ,350, , ,800, ,020, , ,000, ,500 27,525, , ,000, ,500 18,350, , ,650, ,500 15,597, , ,200, ,500 14,680, , ,750, ,500 13,762, , ,750, ,500 9,175, , ,000, ,500 7,340, Totals ,150, ,407, Yearly Average , ,179, $183,500 14,450, Table 2. 3: Estimated Monetary Return on Metro-Sout UAS projet Note the prie-tag derease from 2010 through 2021 depited in Table 2.3. The airraft itself will sell for an average of $308,750/year if you look at the overall time-span from Prodution and development osts for the Metro-Sout are urrently being initially estimated using the DAPCA IV model for lak of a better ost model. The DAPCA IV model was used to estimate: (1) Program Development (RDT&E) osts: Inludes Researh, design, analysis, testing, tooling, engineering, and ertifiation. (2) Prodution osts: Inludes Manufaturing, assembly, materials, quality ontrol, labor, et. April 5, 2007 Page 12 of 72

13 Table 2.4 below shows the breakdown of osts that go into the Metro-Sout UAS over the 1 st five years of prodution ( airframes approx.). Program Cost Predition (Based on DAPCA IV Model) Classifiation Hours Wrap Rate Cost RDT&E: Engineering 85, $7,383,893 Tooling 46, $4,093,843 Development Support $1,254,926 Flight-Testing $1,373,039 Total RDT&E: $14,105,701 Manufaturing: Manufaturing 395, $28,906,066 Quality Control 30, $2,648,271 Mfg. Materials $7,507,978 Total Mfg. Cost: $39,062,315 Payload Flyaway Costs: Loadout $50,000 Avionis $60,000 Engine $10,000 Approx. Prodution ost per airraft: $183,500 Table 2. 4: Cost Analysis for 1st 5 years using DAPCA IV model Team 4 estimates a prodution ost after payload, avionis and engine integration of approximately $183,500 per airframe. Team 4 estimates a market for approximately 950 airframes by 2021 before the Metro-Sout will be retired in favor of newer designs whih would generate sales figures of approximately $120M. Fatoring in the development ost of $14.5M, Team 4 estimates a net return on the produt of around $105M spread over 11 years. (See Table 2.5 below) April 5, 2007 Page 13 of 72

14 Net Sales figures ($ 290,150,000) - Prodution Costs ($ 170,100,000) - Development Costs ($ 14,500,000) = Net Profit ($ 105,550,000) Table 2.5: Estimated net return on produt Team 4 estimates approximately 95 airframes to break even for both prodution and development osts. The following formula was used to arrive at this figure: Development ost + (n x prodution ost) = n x Avg. sale prie (where, n = number of airframes to break even) n = Development Cost / (Avg. Sale prie prodution ost) = $14,500,000 / ($337,000 - $183,500) = 94.5 rounded off to 95 Equation 2.1: Developmental Cost Based on this estimate, the projet would break-even by the early seond quarter of 2012 (i.e, 2 ½ years into prodution assuming ideal onditions). However, if break-even osts were estimated off of the development osts (initial investment) alone, then it would take just 42 airframes to break even for the initial investment of $14.5 Million. Figure 2.8 below harts the yearly profit trend and marks out the break-even point for both prodution and development osts. April 5, 2007 Page 14 of 72

15 Profit vs. Sales by year $120,000,000 $100,000,000 Profit in $ (USD 2007) $80,000,000 $60,000,000 $40,000,000 $20,000,000 Break-Even Point 95 airframes (1 st half, 2012) Series1 $0 -$20,000, Number of Sales (Set-up by year) Figure 2.8: Profit trend vs. number of sales highlighting the break-even point for sales Team 4 will ontinue to refine its ost analysis for the Metro-Sout, and will attempt to find a more aurate ost model for the Metro-Sout UAS. It is interesting to note, however, that other UAVs within the Metro-Sout s weight and size range have similar aquisition osts for example, the Shadow UAV produed for the US military by AAI weighs approximately 400 lbs and has a prie tag of $275,000 it also arries more or less the same type of payload, albeit for military operations. The ost analysis portion of this report is onluded below with a omparison of the osts of the Metro-Sout, and the urrent best selling heliopter for similar operations the Bell Jetranger III. Finanial Baseline 2007 USD Metro-Sout Bell Jet-Ranger Aquisition Cost $350,000 $1,200,000 + payload prie Operating Expenses $10,000/year $20,000/year April 5, 2007 Page 15 of 72

16 3.0 Conept Seletion 3.1 Initial Conepts Figure 3.1: Conept I, Box Wing Conept I, the box wing airraft, has a front wing swept towards the rear and the rear wing swept towards the front. Unlike regular bi-planes, this arrangement does not inlude the poor aerodynami harateristi with vorties from one wing interfering with the other. The main benefit is the signifiant redution of strutural weight in the order of 30%. The sweep also results in good transoni harateristis. Additional tails are not required as the wings provide enough pith and roll ontrol. On the other hand this design is diffiult to manufature. The trimmed maximum lift oeffiient that s equal to the normal wing-tail onfiguration is diffiult to obtain. There an also be exess wetted wing area and interferene drag with many omponent intersetions. [2.1] April 5, 2007 Page 16 of 72

17 Figure 3.2: Conept II, T-Tail Conept II, the T-tail design, features a pusher propeller on the bak of the fuselage with the payload loated at the front. A amera pod hangs from beneath the fuselage fore of the nose gear, whih is part of a triyle landing gear onfiguration. To inrease propeller effiieny by keeping the propeller out of the disturbed air from the horizontal stabilizer, this design utilizes a T-tail. This tail onfiguration plaes the horizontal stabilizer above the inflow to the propeller. The T-tail onfiguration also features a smaller wing aspet ratio when ompared with other designs. A disadvantage to the T-tail involves the extra weight inurred in strengthening the vertial stabilizer to support the horizontal stabilizer. This represents a key detriment to the design of an unmanned aerial vehile for this mission, as the added weight to the very rear of the airraft (where the engine already resides) shifts the enter of gravity quikly aft, dereasing stati margin. Suh a onsideration may null the advantage of reduing the interferene of the horizontal stabilizer with inflow to the propeller. April 5, 2007 Page 17 of 72

18 Figure 3.3: Conept III, Twin Fuselage Conept III, the double fuselage oneptual design, was onsidered very early on, before the payload had been determined. The idea was that the multiple ameras used in the payload ould be spread out between the two fuselages in order to give eah a lear line of sight, but still have a balaned weight. Also, there was the possibility of using the extra spae to retrat the landing gear. One the two payloads were determined to be a single amera, the purpose of the twin fuselage was voided and gave no advantages over the other designs. In fat, the twin fuselage design showed several disadvantages inluding its propeller limitations and exess weight. Figure 3.4: Conept IV, Canard with V-Tail Conept IV, a V-tail arrangement, was onsidered. There are some advantages and some disadvantages to using a V-tail. The V-tail, whih is lighter in weight than a onventional April 5, 2007 Page 18 of 72

19 tail, ontributes to reduing the airraft s wetted area whih in turn redues skin frition drag. Another advantage is that it provides better ground learane than a onventional tail; however, the ground learane would be limited by the pusher propeller engine. The penalty of using V-tail results in ontrol-atuation omplexity. Rudder and elevator ontrol inputs must be blended in a mixer to provide the proper movement of the V-tail ruddervators. In addition, the V-tail auses adverse roll-yaw oupling where the ruddervators produe a rolling moment toward the opposite diretion/turn that the airraft is supposed to do. Thus, as stability and ontrol is one of many issues in designing the UAV, the V-tail oneptual design was deemed undesirable. Figure 3.5: Conept V, Ring Wing Conept V is the most unonventional among the onepts Team 4 seleted for review. The ring wing is a derivative of the box-wing and biplane designs with the wing traing a omplete irle/ellipse around the fuselage. Theoretially, a ring-wing design should provide drag benefits, and added lift, in addition to plaing all the ontrol surfaes on the wing. The problems with this onept arise from stability, design, and manufaturing omplexity required to make it work. These eah add to the ost of the overall projet. April 5, 2007 Page 19 of 72

20 Figure 3.6: Conept VI, Boom Tail Conept VI, the boom tail design was designed to maintain stability of the UAV by plaing the enter of gravity in front of the aerodynami enter. The stability problem assoiated with a push propeller driven airraft stem from the engine weight whih would move the enter of gravity toward the rear end of the airraft. The wing of the airraft would need to be plaed loser to the rear of the airraft ompared to that of a traker propeller or tail mounted pusher propeller driven airraft. Moving the wing bak would redue the moment arm available to the aileron and rudder. A boom tail design overomes the problem by plaing the engine near the enter of the airraft. Center of gravity, in most ases, would fall in front of the aerodynami enter when the payload is plaed at the front of the airraft. The moment arm for the aileron and rudder is thus maintained and the horizontal stabilizer is raised to redue prop wash and sine the propeller is ahead of the tail, there is no interferene from the tail to redue propeller effiieny. The boom tail, however, would inrease the weight of the airraft in order to reate a sturdy tail struture. April 5, 2007 Page 20 of 72

21 Figure 3.7: Conept VII, Canard Pusher-prop w/ Vertial Tail Conept VII depits a pusher-prop anard airraft with a single vertial tail, and fixed triyle landing gear. This is, for the most part, a onventional anard-type airraft reminisent of most fighter airraft designs that inorporate anards, and differing from most of the Burt Rutan designs in terms of not having vertial tails at the wingtips. Also, the main wings have not been swept beause the payload weight in the nose provides suffiient ballast to maintain the forward enter of gravity within ontrollable limits so that the engine weight annot tip the airplane over during rotation for takeoff or landing. This negated the need for wing sweep to move the engine loser to the airraft enter of gravity. The anard onfiguration also provides desirable stability harateristis. When designed orretly, the anard will stall before the main wing, pushing the nose down and preventing the wing from stalling. April 5, 2007 Page 21 of 72

22 Figure 3.8: Conept VIII, Cruiform Conept VIII is a onventional airraft design. It has a vertial tail with horizontal stabilizers. This partiular design also inorporates a pusher propeller engine and fixed triyle landing gear. This is the general design for most general aviation airraft suh as the Cessna 172. There are very few disadvantages to this partiular design but as mentioned above there are many more advantages to designs that allow for higher lift. High lift is extremely important for the urrent mission of the Metro-Sout as it is a small, light airraft that will loiter for approximately 5 hours. 3.2 Pugh s Method After oneptualizing several initial design possibilities for the airraft, the team needed to selet the most appropriate onept for the mission. To aid in this seletion, the team utilized a tool alled Pugh s method of onept seletion [3.1]. The goal of this proess involves omparing the initial designs generated based on the requirements from the House of Quality developed for system requirement analysis, developing new onepts from the positive aspets of the initial designs, and finally deiding on the best design for the partiular mission. A design group rarely selets an initial onept without modifiation for the final design, as was the ase for the team in harge of designing the Metro-Sout. Pugh s method relies on a matrix of riteria and design onepts as a visual means of omparing aspets of andidate onfigurations. An example of suh a matrix is found in Table 3.1. CONCEPTS CRITERIA A B C D Table 3.1: Pugh s Method Template The design team initially used riteria diretly orresponding to the engineering expetations that omprised the olumn headings of its QFD House of Quality. This method, however, led to many awkward if not useless omparisons between onepts due to April 5, 2007 Page 22 of 72

23 lak of prior knowledge. For example, one riterion used was L/D (lift-to-drag) ratio. Comparing designs based on this harateristi proved nearly impossible beause the team had no way of knowing what the final L/D ratio would be for eah initial onept. Further development of the riteria led to ideas that ould be easily evaluated with the team s base knowledge, suh as manufaturing ost. One the riteria had been seleted, the team then seleted the original onept it felt best met the riteria and mission at the time. This beame the datum for omparison in the matrix. Group members then ompared eah design to the datum based on the seleted riteria using the following notation entered into the orresponding system: + : onept meets riterion better than datum S : onept not learly better or worse at meeting the riterion than the datum - : onept meets riterion worse than datum. Table 3.2 displays the Pugh matrix from table 3.1 after a datum is hosen and ompared with the other onepts. CONCEPTS CRITERIA A + DATUM - + B S DATUM - + C S DATUM + + D - DATUM S S Table 3.2: Pugh s Method Example In table 3.2, onept two (2) represents the datum. The matrix shows that onept one (1) is more profiient at meeting riteria A than onept two, but not as profiient at meeting riteria D. This matrix seems to indiate that onept four (4) possesses several advantages over the datum and, in effet, might omplete the mission more effetively. Pugh s method of seletion, however, urges a designer to determine whih features of a design are most effetive, not neessarily the entire design. The team also added the riteria of design life, April 5, 2007 Page 23 of 72

24 manufaturing ost, and operating ost, onsiderations that were not present in the House of Quality. Based on the first iteration of the Pugh matrix, the team developed several new onepts by eliminating adverse aspets and adding advantageous features to the first designs. They then ran the matrix with these new onepts and a new datum. The finished produt represented the result of three iterations of the matrix, eah with new onepts and new datum designs that produed a synthesis of onepts that best satisfied all riteria. Setion 3.1 desribes the initial design onepts drawn by eah team member. After developing the proper riteria, the team onduted three iterations of the Pugh matrix using the initial onepts for omparison. Table 3.3 displays the results of the first iteration of the method. For this initial omparison the team seleted the boom tail design as the datum onsidering its lak of interferene with pusher propeller inflow while allowing room for substantial horizontal and vertial stabilizer surfaes. Team members ompared eah design to the datum in terms of every riterion. As an example of the omparison proess, group members onsidered eah design in terms of the riteria of lift and drag (rows ten and eleven in table 3.3). As a matter of referene, designers aimed to minimize drag and maximize lift. Row ten shows that, aside from the T- tail design, no design met the riterion of minimum drag to any greater or lesser extent than the datum boom tail design. The team surmised that the T-tail might reate a large amount of drag at moderate to high angles of attak where the horizontal stabilizer would be plaed diretly in the turbulent outflow from the propeller. Hene, this design warranted a - evaluation in this ategory. In terms of lift, the Metro-Sout team felt that the box wing design would produe more lift than the boom tail, whih involved a standard wing. This deision hinged on the knowledge that well-designed box wings reate more lift than standard wing designs. This is the key advantage to a box wing, hene it earns a + evaluation when ompared to the datum. The T-tail reeived a - in the lift ategory also. This is due to its lower aspet ratio than the April 5, 2007 Page 24 of 72

25 boom tail design. No other design qualified for an advantage or disadvantage in the ategory of lift. Table 3.3: First Iteration of Pugh s Method After the team ompleted its evaluation of every onept in eah of the sixteen design riteria, the team summed the positive and negative evaluation marks for eah design. This ation did not give eah design a sore but rather helped the group evaluate what was poor and advantageous for eah design. After areful evaluation of eah of these aspets, the anard design beame the new datum. While onerns about stability of the anard design existed, it ompared well to the other ideas in terms of overall weight (row 1), strutural loading apaity (row 6), and design life (row 14). It also met all other ategories equally as well as the boom tail design. Using several hybrid onepts as well as some modified original ideas from the first iteration for omparison with the anard design as the datum, the team performed a seond and third iteration of the Pugh matrix. Table 3.4 shows the third iteration, whih represents the final design seletion of the anard onfiguration. This iteration indiated that no other onept met the design riteria as effiiently as the anard design. The team deemed some April 5, 2007 Page 25 of 72

26 areas where the anard seemed to be laking (stability, manufaturing ost, operating ost) reoverable through design. For example, the anard design had potential for being quite unstable without areful onsideration, but the team deided it ould overome this disadvantage with proper layout and aerodynami design. Business ase and market analysis also indiated that advantages in design ould generate enough additional revenue to null the advantages of the other designs in the ategories of manufaturing and operational osts. Table 3.4: Third Iteration Table 3.4 shows that the team inluded a net value for determining if one design had an overall advantage for this iteration. Don Clausing advises against reating a sore for any given onept when using Pugh s method [3.1], but the team felt that this iteration served mainly to ompare what it thought would be its best design to some previous ideas. In this way it ould be ertain its ore onept would be the best to omplete the mission. The net April 5, 2007 Page 26 of 72

27 results seemed to support the idea that the team should work toward a anard design, as no onept possessed a net advantage over it. While the team s Pugh analysis produed the anard design as the potential best onept for the Metro-Sout design, it leaves open room for design modifiations and improvements. The team may update the matrix at some point during the design should new information beome available that may give another design aspet an advantage in a given ategory. Again, the end goal of the Pugh method involves produing the best overall onept, and in a hanging market and design environment, the requirements of the airraft may hange over the ourse of its development. 3.3 Seleted Conept Team 4 has seleted the single vertial tail anard pusher-prop airraft onept (onept VII) to be the most viable solution for the mission engineering requirements, based on the results from the use of Pugh s method of onept seletion. Shown below is a walk-around view of the seleted onept, detailing the internal and external layout of the Metro-Sout. (Please note that the Metro-Sout will have a fixed triyle landing gear set-up, but that it is not depited on the CAD walk-around view below as the gear itself has not been designed yet. Also, the mounts for the internal omponents will not be attahed to the fuselage skin in the final design, and only appear to be so as the final oneptual strutural layout has not been finalized by Team 4 yet.) April 5, 2007 Page 27 of 72

28 Figure 3.1: CAD walk-around image of the Metro-Sout onept - Note the Canards and Pusher Prop The external layout of the Metro-Sout design is detailed on page 33 and the internal layout is detailed on page 36. Team 4 has listed the speifi plus-points and hallenges assoiated with using a pusherprop and anards in the following setions. Use of a Pusher Prop: Highlights and Disussion: Team 4 has eleted to use a pusher-prop on the Metro-Sout design for the following reasons: (1) Visibility: A trator prop design would plae the rotating propeller diretly in the view-frame of the amera system and inhibit the visibility of targets that the amera was trying to trak. A pusher-prop design plaes the propeller at the rear of the fuselage and out of the way of both the pilot and payload ameras. (2) Slipstream Effet: The prop-wash oming off of a forward mounted propeller auses a number of undesirable effets on the wings, tail and ontrol surfaes. This is April 5, 2007 Page 28 of 72

29 espeially notieable for anards where sudden hanges in the engine throttle setting have been known to ause stalls, violent pithing motions, and even fatal rashes due to the slipstream effet. Changes in the throttle setting ause a notieable differene in the veloity of the propeller slipstream that affets the ontrol surfaes, and in turn auses the airraft to experiene distintly observable pith and yaw hanges during power hanges. This effet is summarized below in figure 3.2. Figure 3.1: Slipstream effet on a Cessna 172 (Image Soure: (3) Exposed Area & Resultant Drag: A mid-fuselage mounted engine-propeller arrangement suh as observed in airraft like the Lake Renegade (pitured below) adds a large amount of exposed area to the free-stream whih tremendously inreases drag. Suh an arrangement is useful primarily for keeping the prop ompletely out of reah of the ground or in the ase of sea-planes, out of danger of hitting water. A pusher-prop is relatively effiient in the aspet of exposed area and drag as the airflow an be made to remain relatively laminar over the fuselage leading into the propeller and the usually flat non-streamline surfaes of the engine housing are not exposed to the free-stream. Additionally, a fuselage-mounted pusher prop redues the wetted area of the airraft by shortening the fuselage. [2.1] Team 4 will size the landing gear on the Metro-Sout to ensure adequate ground learane for landing and take-off. April 5, 2007 Page 29 of 72

30 Figure 3.2: Mid-Fuselage Mounted Engine on a Lake Renegade (Image Soure: (4) Redued Airraft Skin Frition Drag: A pusher prop design allows the fuselage, wing, anards and tail to fly in a region of undisturbed air. Not flying in the propwash reated by a trator prop onfiguration redues the airraft skin frition oeffiient, and so, the overall drag of the airraft. [2.1] (5) Redued Cabin/Payload Bay Noise and Vibrations: In the urrent onfiguration of the Metro-Sout, the payload is at the opposite end of the fuselage from the engine this serves the benefit of redued vibrations and engine noise being felt in the payload ompartment. The ameras and avionis are expensive, sensitive equipment, and a pusher-prop design goes a long way toward inreasing the life-span and dereasing the maintenane osts of the payload equipment. [2.1] (6) Canard-Pusher Combination Advantage: Canard airraft are usually designed with a pusher-prop beause of the additional benefit of having a shorter tail arm as ompared to an aft tail. [2.1] Naturally, as with any type of design, there are a unique set of hallenges assoiated with a pusher-prop design. These are listed below: (1) Ground Clearane: Most notably, pusher-prop designs need longer landing gear to ensure adequate ground learane for taxi, take-off and landing. Soft-field and rough field landings are a speial hallenge for pusher-props as dirt, roks or turf an get kiked into the propeller blades. It is Team 4 s deision that the Metro-Sout annot and does not need to, given the nature of its ustomers, be ertified for types of operations requiring rough field landings where a risk of FOD (foreign objet damage) into the propeller blades an our. [2.1] April 5, 2007 Page 30 of 72

31 (2) Redued Effiieny: In pusher-prop designs, the propeller ats in a region of disturbed flow oming off of the fuselage, wings and anards, and so, has a ertain amount of redued effiieny. [2.1] For the Metro-Sout design, an attempt will be made to keep the flow oming off the fuselage relatively laminar. However, given the relative parasite drag-savings gained by not having prop-wash at on the fuselage, anards, and wings, the redution in propeller effiieny an be negated with the right design. (3) Engine Cooling: Due to the loation of the engine, engine overheating is an issue that a number of anard pusher-prop designs have had to overome. Team 4 will look into positioning air inlets on the fuselage speifially for engine ooling in loations of uninterrupted airflow where attitude hanges will not result in signifiant flow-rate hanges. (4) Aural Signature: Pusher-prop designs have a notoriety for being noisy ompared to onventional trator-prop designs. This is primarily due to the engine exhaust flowing through the propeller blades generating a harateristi whine. Additionally, airflow shearing off disontinuities in the fuselage, ontrol surfaes, turbulent air oming off the wings an ontribute to an inreased noise signature as it passes through the propeller. Given that the Metro-Sout will be operated over metropolitan areas, reduing propeller noise is a legitimate design objetive. Team 4 will look into ways to rediret engine exhaust away from the propeller blades suh as through exhaust pipes that vent out the trailing edge of the vertial tail or the wings. Additionally, every attempt will be made to keep flow laminar over the fuselage by minimizing protrusions or disontinuities in the fuselage. Use of Canards: Highlights and Disussion On an intuitive level, from an engineering standpoint, Team 4 an subjetively say that a Canard airraft design, if done right, an give the Metro-Sout produt advantages in terms of maneuverability, safety, fuel-effiieny, and weight & drag-savings that far outweigh the potential disadvantages in airframe omplexity, and inrease in time and ost for design, analysis, optimization and flight-testing. To further elaborate on this, table 3.5 below highlights the following items: April 5, 2007 Page 31 of 72

32 (1) Positives and negatives that Team 4 has established for a Canard airraft as it relates to the speifi mission apabilities required for the Metro-Sout Produt (2) Some historially observed trends for advantages and disadvantages of Canard type airraft Advantages Typial Canard Airraft Charateristis Disadvantages/Challenges (1) Good Stall Charateristis/Can prevent stall (2) Pusher-Prop Pakaging/Assembly an be muh simplified (3) Canard lift an be made to ompliment main wing lift (4) Fuselage supported in two plaes Some weight savings (5) Sometimes more useful range of.g. (6) Can have added Maneuverability (7) Can have added fuel-effiieny (1) Canard Sizing is highly ritial small hanges an affet performane (2) Downwash/Upwash effets on the main wing from the anard (3) Suseptible to Deep Stall if pilot/operator over-maneuvers airframe (4) Small moment arm of Canard leading to larger anard area (5) Added time for wind-tunnel analysis and flight-testing Table 3.5: Observed Advantages and Disadvantages of using Canards in Airraft Design [2.1], [2.3] As mentioned earlier, how muh eah of these potential advantages and disadvantages impat airraft performane harateristis are dependent uniquely upon the exat nature of the Metro-Sout design. With a properly designed airraft that inorporates design elements that build upon the lessons learnt from available literature on previous Canard designs, Team 4 believes that the Metro-sout design an be quite suessful and virtually problemfree. Team 4 has already onduted preliminary aerodynami and stability analysis into the Metro-Sout anard onept. Current estimates indiate that a high stati-margin on the order of 12-15% an be ahieved using anards with a span of feet, and a hord length of 1 foot. Currently, the anards omplement the lift generated by the main wing to ahieve greater effiieny. The aim is to ounter the redution in lift on the main wing by putting in some lift on the anard. Also, the plaement of the anards and the main wings April 5, 2007 Page 32 of 72

33 redue redundany in strutural load bearing elements. For instane, the strutural members holding the anard to the fuselage also partially bear the weight of the payload and the front of the airraft. Similarly, the engine, the vertial tail and the main wing share a rigid ommon root support-struture that onnets them to the rest of the fuselage. Team 4 will ontinue to evaluate the sizing of the anards on the Metro-Sout to ahieve the most effiient balane between performane, stability and fuel-effiieny for the airraft. In addition, Team 4, as a result of its experiene thus far in the aerospae ulture, has noted the inherent ulture of resistane in both ustomers and manufaturers in the ommerial airraft market to the mass-prodution of non-onventionally designed airraft. Team 4 onsiders that the use of anard tehnology in airraft has reahed a suffiient level of tehnology maturity to enable both ustomers and the industry to aept that the performane, reliability, manufaturability and finanial viability are omparable to that of a onventional airraft design that would perform the same funtion. A lot of this level of tehnology maturity, at least in the modern era, an be attributed to the work of noted airraft engineer and designer Burt Rutan. The use of anards in Rutan s designs inorporates elements of design simpliity, ost-effetivity, safety & maneuverability, and high performane. The Rutan-designed Long-EZ and Proteus airraft (pitured below) have ahieved high endurane times far beyond other airraft in their respetive lasses as a result of their effetive use of anards. [2.3] Figure 3.3: Proteus Airraft (Saled Composites/Northrop Grumman), (Image soure: NASA) April 5, 2007 Page 33 of 72

34 Team 4 has noted that manufaturing and prodution times do not seem to be signifiantly affeted either positively or negatively by going with a Canard type design. Delays and problems with the manufaturing proess in historial examples of the development of some anard airraft ould be attributed to inaurate tasking, onfusion among manufaturing personnel resulting from misommuniation of proedures, and in one unique ase involving a Japanese/German fighter design- the end of World War II. But learly, these are problems that are not intrinsi to the use of anards. 3.4 External Layout Figure 3.4: CAD 4-view of the External Layout of the Metro-Sout The Metro-Sout is sized based primarily on the following requirements: April 5, 2007 Page 34 of 72

35 (1) To allow all internal equipment suh as payload, engine, avionis, and struture to be mounted safely and with ample room to allow easy aess for maintenane, and freedom of interferene from neighboring equipment. (2) To allow the positioning of the wings, anards, vertial tail and ontrol surfaes to ahieve the desired levels of stability, maneuverability, and ontrollability in all flight onditions. Figures 3.5 above and 3.6 below show some initial external dimensioning on the Metro- Sout. Figure 3.5: CAD Sketh of the Metro-Sout with additional dimensioning and internal utaway The main wing, as with most anard airraft is plaed toward the rear of the fuselage in order to position the enter of gravity in an optimum loation. The engine and propeller are loated diretly behind the main wing box. In addition, the vertial tail is plaed diretly behind the main wing box. An additional benefit of this positioning is that the engine, the April 5, 2007 Page 35 of 72

36 vertial tail, and the main wing share a rigid ommon root support-struture that onnets them to the rest of the fuselage. The main wings are not swept bak beause the payload weight in the nose provides suffiient ballast to maintain the.g. within ontrollable limits so that the engine weight annot tip the airplane over during rotation for takeoff or landing. This negated the need for wing sweep to move the engine loser to the airraft enter of gravity. Currently the main wing span is 26 feet, with a root hord of approx. 5 feet, and a tip hord of 2.5 feet. The sizing of the wing is based on initial aerodynami analysis that Team 4 onduted, whih is detailed in setion 5 of this report. Based on the requirements for stability and ontrollability, the anard was plaed at a distane of 4 feet from the nose of the fuselage. The anard span is 14 feet from tip to tip. It is based on a retangular planform with a hord of approximately 1 foot. 3.5 Internal Layout The fuselage of the airraft was designed in three separate parts, eah with a different driving fore behind the design. The position of the omponents of the airraft (payload, anard, transmitter, fuel tank, wing, tail and engine) had already been roughly hosen, so the fuselage was split up aording to whih omponents it would need to house. The front setion houses the payload and the anard and requires a bubble of glass on the belly so that the ameras will not add to drag but still be able to funtion. The mid setion houses the fuel tank and the transmitter and the last setion houses the wing box and tail struture. The front of the fuselage needs to be wide enough to aommodate a 24 inh x 24 inh ube whih represents the area in whih the largest amera an move. Giving the ube a 2 inh buffer from any point on the fuselage ensures that any vibrations will not ause the amera to hit the glass bubble. The 8 inhes of distane between the top of the amera mount and the top of the fuselage allows suffiient spae for both mounting strutures and wiring. The anards are also plaed in this setion of the fuselage, loated behind the payload. The aerodynami enter of the anard is 3.7 feet behind the nose of the airraft. April 5, 2007 Page 36 of 72

37 The size of the fuel tank was the driving fore behind the design of the next setion of the fuselage, between the bubble and the wing. The fuel tank needs to be loated outside and in front of the wings beause of the wing box, the small thikness of the wing, and in order to move the enter of gravity forward. A short and wide tank was hosen so that as the fuel level lowers and the fuel has more freedom of movement, it will not hange the.g. signifiantly during maneuvering. Also, the fuselage needed to be as skinny as possible in order to redue surfae drag. The final tank size was inhes x 26 inhes x 14inhes. This fits into the midsetion of the fuselage with a 7 inh learane on eah side and a 5 inh learane on the top and bottom. This is enough spae for strutures to mount the fuel tank. The mirowave transmitter was also plaed in this setion, just forward of the fuel tank. This 12 inh x 9inh x 4inh box weighs approximately 11 pounds and will share a mounting platform with the autopilot ontroller, whih is of negligible size and weight. The last setion of the fuselage, whih will extend from behind the fuel tank to the trailing edge of the tail will house the wing box and tail struture. Neither of these has been designed yet but they are antiipated to require that the size of the fuselage will be larger than the fuselage surrounding the fuel tank and mirowave transmitter. Though it is modeled as an inboard engine in the Catia model, a deision has been made to mount the engine outside of the fuselage in order to use the air to ool it. Figure 3.7 is a wire view of the Metro-Sout. This view shows the dimensions of the payload pay and the fuselage. The side view of the airraft shows the distane from the nose to the payload, anard, transmitter, fuel tank, wing and the total length of the airraft. Note: In the final design the transmitter will be moved down to make room for the anard. April 5, 2007 Page 37 of 72

38 Figure 3.7 Inboard Layout The team is not at the point in the design where internal fuselage strutures have been designed. That is the reason that the Catia model shows the representations of the omponent mountings as being attahed to the skin of the fuselage. In future steps, ribs will be plaed to give the fuselage the appropriate strength, as well as provide the struture needed to support the omponent mounts. Also being developed is the wing box and tail struture. 3.6 Payload Integration Different payloads are required for the different missions designed for the law enforement agenies and the news agenies. These payloads were reated based on the onept of operations desribed above and the different ustomer attributes requested. Table 3.6 lays out exatly what piees of equipment are ompiled to make up the payload pakage that will be sold to law enforement. April 5, 2007 Page 38 of 72

39 Pakage Payload Weight(lb) Dimensions(ft) Radar gun x 0.19 x 0.45 Polie pakage Camera gimbal x 1.25 ThermaCAM SC mounted on gimbal Sony DSR-PD150 (video am) 3.1 mounted on gimbal Canon powershot S3 IS ( still amera) 0.9 mounted on gimbal Canon lens f/ II USM 0.7 on amera Autopilot x.17 x.14 Total weight 64 Table 3.6:Payload Pakage for Law Enforement[3.2],[3.3],[3.4],[3.5],[3.6],[3.7],[3.8] During highway patrols for speeding vehiles, the Metro-Sout will fly at the speed limit set for the highway and any ar moving faster than the UAV will trigger the radar gun to reord the exat speed of the speeding vehile. This in turns triggers the still amera to snap a piture of the vehile s liense plate. Law enforement offiers an then issue the violators a tiket and mail it to them. In searh and resue operations, the video amera, infrared and still amera will work in tandem with one another. If the infrared amera detets a possible target, the video and still ameras will be used to positively identify the target. These images are then sent bak via live feeds through a transmitter. The radar gun has an auray of 1.25 miles per hour and is able to measure a target speed moving in a o-diretion takes plae if the speed differene is varied from 2.5 up to 62 miles per hour. It is also not important where the target is loated in front of UAV is or behind the UAV, the UAV athes up with target or the UAV is left behind target - in any ase the orret evaluation of a target speed is guaranteed by Semion. This lightweight radar gun of less than 1 pound will be plaed on the UAV and feed bak information on traffi violators. April 5, 2007 Page 39 of 72

40 The Camera gimbal has a 4-axis gyro stabilized video system. It is able to rotate 360 degrees and trak stationary and moving targets from up to 3000 feeht. [3.3] The amera operator an ontrol the gimbals system whih then transmits the video feeds and still images through mirowaves bak to the ground. The ThermaCAM is an infrared amera oupled with its software an provide live feeds of an extensive temperature range. It is able to measure extremely small and distant targets with great auray (±1%) and high resolution. [3.4] This will assist law enforement in riminal pursuit during the day and even at night. In the event that the suspet is hidden from the regular view of a regular amera, the infrared an still detet the heat signature of the suspet. This infrared amera is also relatively light at seven pounds, providing an additional apability whih aids the apture of riminal suspets. Most importantly, a video amera is also mounted on the gimbal. The Sony DSR-PD150 has a build in image stabilizer with a 12 X optial and 48X digital zoom. [3.5] This amera will be able to provide lose up aerial videos from the air. This amera is also very light weighing only 3.1 pounds. A high resolution still amera is also essential for law enforement agenies. The Canon powershot S3 IS will be equipped with a millimeter lens that will be able to take high resolution still images of liense plate numbers from up to a distane of 3000 feet. [3.7] This amera an also be used by law enforement to take high quality pitures of evidene against fleeing suspets. The autopilot s software for the UAV is apable of flying at a Maximum Altitude of 16,000 feet above sea level and a maximum airspeed of 150 miles per hour. It omes with a transmitter to broadast airspeed, pressure and temperature to the ground in ompliane with FAA regulations. April 5, 2007 Page 40 of 72

41 The payloads for the news stations are as follows: Pakage Payload Weight(lb) Dimensions(ft) Cineflex V X 1.63 X 1.63 News Station/filming pakage Autopilot X 0.17 X 0.14 Total weight Table 3.7:Payload Pakage for News Ageny [3.8],[3.9] The amera for the news station allows for high definition live video feeds. This amera weighs more than the entire payload for the law enforement as it is a high definition filming amera mounted on a gimbal, but the 67lbs inludes the weight of the gimbal. The Cineflex amera is also urrently mounted on heliopters and also used for filming movies. The built in wide angle view and infrared ameras allow for filming at all times of the day. Lastly, the Cineflex also has a 25X zoom enabling aerial footages to be filmed from up to 3000 feet. [3.9] 4.0 Constraint Analysis 4.1 Constraints The performane analysis, in most ases, answers the question of whether a partiular airraft design will meet a ustomer s needs. The proess of onstraint analysis is to narrow down the hoies of the many interrelated variables to ontrol and make hoies to whih to design an airraft suh that it will have the desired performane apabilities. Constraint analysis provides ranges of values for an airraft onept s take-off wing loading and takeoff power loading, whih allow the design to meet speifi performane requirements. The onstraint analysis is based on a modifiation on equation 4.1 for speifi exess power T D 1 dh dv = W W V dt g dt In equation 4.1, T/W is the thrust to weight ratio, D is drag, V is veloity, dh/ht is the altitude derivative and dv/dt is the veloity derivative. By substituting equations 4.2, 4.3, 4.4 and 4.5 into equation 4.1, the new onstraint equation is stated in equation 4.6. April 5, 2007 Page 41 of 72

42 T W β q C = α β W ( TO T = αt SL 4.2 W = βw TO 4.3 L nw C L = = qs qs CL D = CD qs = ( CD0 + ) qs πa Re nβ WTO 1 dh dv ) πa Re q S V dt g dt S 4.6 SL D 1 TO In equation 4.2, α is the thrust lapse ratio whih depends on the density ratio ρ equation 4.3, β is the weight fration for a given onstraint. This fuel fration is neessary beause the weight loss from the fuel has to be taken into onsideration at every moment throughout the flight. Equation 4.4 is the equation for the lift oeffiient. Equation 4.5 is the drag equation based on the lift oeffiient found in equation 4.4. Equation 4.6 is the newly defined power equation for take-off weight. [2.1] ρ SL. In 4.2 Takeoff Constraint While equation 4.6 models in-flight performane, the takeoff onstraint requires a different equation. AssumingV = 1. 2V, equations 4.7, 4.8, and 4.9 are written below. TO stall V TO 2W TO = ρsc L max t TO = 1.2W 2W 2 { g[ T D μ( W L)]} ρsc 4.8 TO L max s TO = 1 2 at 2 TO = ρsc L max 1.44W 2 TO g[ T D μ( W TO L)] 4.9 Rewriting equation 4.6 using equations 4.7, 4.8 and 4.9, equation 4.11 is the new power equation in terms of power loading, equation 4.10, and wing loading. April 5, 2007 Page 42 of 72

43 P T SL VTO = 4.10 W TO WTO 550η p 2 TSL 1.44β WTO CD0q = + + μ 4.11 W αρc max gs S W TO L TO β ( TO ) S The unit of power in the above equation is horsepower. These equations also assume that lift is approximately zero prior to rotation. [2.1] 4.3 Sustained Turn Constraint Maximizing thrust loading and lift to drag ratio (L/D) maximizes the load fator in a sustained turn. At max L/D, the oeffiient of drag isc D0, therefore deriving equation W q = πa ReC D S n Equation 4.12 is the wing loading equation for the max range and max propeller loiter for a propeller airraft. This equation proves that as weight redues due to fuel burned, the wing loading also dereases during ruise. Optimizing ruise effiieny while wing loading is dereasing requires the redution of the dynami pressure by the same perent as seen in equation The onept of max L/D and the above wing loading equation yields equation 4.13; the available thrust equation. [2.1] T W = qcd W S 0 2 W n + S qπae Landing Constraint The landing onstraint determines the maximum value of wing loading of the UAV. The maximum wing loading bounded by the landing onstraint is alulated from the landing onstraint equation below. W S In equations 4.14 d land is the landing distane, land W β is the landing weight fration ( land ). W d land ρcl max gμ = β μ is the frition oeffiient when landing, o April 5, 2007 Page 43 of 72

44 4.5 Constraint Results By running MATLAB ode developed by team members, the group determined a design point for power loading and wing loading. Figure 4.1 shows this design point in terms of a speified power loading and wing loading value. [2.1] Power Loading [hp/lb] Power Loading Constraint Analysis Loiter Turn Constraint with Load Fator =3 Max Speed Turn Constraint with Load Fator =1.5 Takeoff onstraint for takeoff distane=250ft Loiter Steady Flight Constraint for 73 ft/s Max Speed Steady Flight Constraint for 176 ft/s Aleration Constraint for 10 ft/s 2 Landing Constraint for Landing distane = 100 ft Design Point Wing Loading [lb/ft 2 ] Figure 4.1: Power Loading Constraint Analysis This point represents a ratio of power loading to wing loading that best blends the loiter turn, max speed turn, and takeoff onstraints. It does not neessarily optimize any of these onstraints, but rather suggests a point where hanging one onstraint would adversely affet another. Table 4.1 below, shows the optimum design point. Power Loading hp/lb Wing Loading 7.8 lb/ ft 2 Table 4.1: Power Loading and Wing Loading Data The driving onstraints were the aeleration and the max-speed turn load fator onstraints. The driving onstraints determined the power loading and wing loading of the UAV. The non-driving onstraints are also shown in figure 4.1 to demonstrate the apability of the UAV. April 5, 2007 Page 44 of 72

45 Mission Requirement Value Ahieved Max-speed Turn 1.5g 1.5g Load Fator Aeleration for 10 ft/s 2 10 ft/s 2 loiter to max speed Take-off distane 1500 ft 250 ft Loiter-speed Load 2.5g 3.0g Fator Landing Distane 1500 ft 100 ft Table 4. 1 : Performane Capability 4.6 Sizing Sizing is the proess to determine how large the airraft must be to arry enough fuel and payload to be able to loiter for up to five hours and to take surveillane for new agenies and law enforement. A rude estimate of the maximum L/D is obtained. Speifi fuel onsumption is dependent on the engine hosen for the UAV whih is disussed in setion 6. SFC (C hpb) was taken to be 0.52 and 0.56 lb/hp/hr for ruise and loiter respetively [5.1]. C D0 was obtained from the wetted area alulated for the estimated shape of the UAV. Sine empty weight is alulated using a guess of the takeoff weight, it is neessary to iterate towards a solution. The initial empty weight fration was obtained from regression analysis on historial similar UAVs. The empty weight is an estimation of ombination of all omponent weight unertainties. Equation 4.15 is the atual equation used to determine the empty weight fration of the Metro-Sout. W W e W 0 AR Powerloading Wingloading VMAX = 4.15 The team has developed a MATLAB program to alulate the total takeoff weight. Value Inputted Power loading hp/lb Wing loading 7.8 lb/ft 2 SFC for ruise 0.52 lb/hp.hr SFC for loiter 0.52 lb/hp.hr AR 13 April 5, 2007 Page 45 of 72

46 L/D 13 C D η 0.80 p Oswald s effiieny 0.8 Table.4.2 : Inputs of Sizing Program 4.7 Fuel Fration for Cruise and Loiter The gross weight equation is based on the fuel fration and the empty weight fration. Both of these equations are based on L/D. The L/D equation shown in equation 4.16, is derived on the premise thatc D0 is.0239, AR is 13, and e is This equation is also under the assumption that V loiter is 73ft/s and V ruise is 176 ft/s. L D 1 = qcd0 W + ( W / S) S 1 qπa Re The fuel fration and weight equations derived from the Breguet equation for ruise and 4.16 loiter, used to find the gross weight, are shown in equations 4.17 and 4.18 respetively. W W i RCbhp = exp 550η p ( L / i 1 D) 4.17 W EVC i bhp = exp 4.18 Wi 1 550η p ( L / D) In these equations R is range, E is endurane, C bhp is the speifi fuel onsumption for propeller airraft. η p is the propeller effiieny. The i index in the above equations is the segment number. In any given flight there are multiple segments. For example, in a normal flight, the loiter segment would by i=4 after the take-off (i=1), limb (i=2), and ruise (i=3). The airraft weight is alulated throughout the mission. For eah segment the airraft weight is redue by fuel burned. Total fuel burned is alulated throughout the mission and found by summing the weight frations from eah flight segment in equation April 5, 2007 Page 46 of 72

47 = x W fuel W fi 4.19 i In equation 4.19, 6% of extra fuel is added for landing, takeoff, taxi and reserve. Equation 4.20 is the takeoff weight equation. W W = e 0 W pay W fuel W0 W0 This equation is a summation of the different weights alulated from the various fuel frations. [5.9] Using data olleted from the above analysis, table 4.4 is a ompilation of the initial sizing results. Total Airraft Takeoff weight 603 lb Fuel Weight 116 lb Payload Weight 63.1 lbs Airraft Inert Weight 424 lb Power Required 31 HP Wing Area 77 ft 2 Table 4.4: Sizing Data 4.8 Carpet Plot The arpet plot shows a diret relationship between gross take-off weight, and a range of wing loading and power loading. It also provides estimates of the gross take-off weight with variants in wing loading and power loading. The arpet plot is generated by inputting a range of wing loading and power loading values into the sizing MATLAB ode developed by the team. April 5, 2007 Page 47 of 72

48 750 Carpet plot with onstrants Powerloading= Take-off Weight Design Point Take-off Constraint Aleration Constraint 550 Powerloading= Aspet Ratio Analysis Wing Loading [lb/s 2 ] Figure 4.2 Carpet Plot By generating arpet plot for a range of aspet ratios, and taking the lowest talk-off gross weight for eah aspet ratio, a plot of the take-off gross weight versus aspet ratio is shown below. 625 Take-off Weight vs. Aspet Ratio 620 Take-off Weight [lb] Design Point Aspet Ratio Figure 4.3 Take-off Weight versus Aspet Ratio April 5, 2007 Page 48 of 72

49 The plot shows the optimal aspet ratio at 13. Currently the team does not feel omfortable with this high aspet ratio and is working with the weight equations to get bak to a more suitable number. For the purposes of initial sizing and onept seletion, Team 4 seleted an aspet ratio of 6.8 for the Metro-Sout main wing. The rationale behind this initial aspet ratio seletion was that historially, general aviation and homebuilt airraft have an average aspet ratio of 6.8. [Raymer, 2.1] Team 4 felt that the Metro-Sout UAV would most likely fit somewhere between those two ategories of airraft. However, in future analysis of fuel-savings, weight-savings, and strutural analysis, Team 4 antiipates that the final aspet ratio seleted will be somewhere in the 9-12 range. 5.0 Aerodynami Analysis 5.1 Airfoil Seletion In seleting an airfoil, the airraft design requirements must be found, suh as how it should perform and how it should handle. In general, a higher setion oeffiient of lift ( l ) auses in a higher setion oeffiient of moment ( m ) during ruise. [1] As a result of this pithing moment, the anard must, in turn, provide neessary lift to balane the nose down effet whih in turn leads to a higher trim drag. From the ustomers and mission requirements, a loiter speed of 50miles per hour and maximum speed of 120 miles per hour were below 130 miles per hour at whih NACA airfoils are proven to work. Beause of this, the analysis was done using 4-digit NACA airfoils. [5.1] The 4-digit series was also hosen due to its small enter of pressure movement aross a large speed range. [5.1] The main riteria of the airfoil seletion were a high l0 and a high lmax with l over oeffiient of drag ( di ) and m as seondary requirements. This was beause of the need for a high l0 to ruise effiiently without requiring a larger planform area at the required loiter veloity of 50 miles per hour. A high lmax would provide a higher wing loading for a shorter takeoff distane and for a better sustained turn rate. [5.2] A low pithing moment lose to zero about the aerodynami enter would lower trim drag indued by the anard. April 5, 2007 Page 49 of 72

50 This, however, an be adjusted by adding ounter weights suh as payloads or varying the loation of the fuel tank in the fuselage to minimize this effet. The analysis was done with varying amber, while the loation of the amber from the leading edge remained onstant at 40% hord length and the thikness remained at 12% hord length. As the loation of the amber from the leading edge inreased, l inreased, thus the loation of the amber from the leading edge was piked to be as far bak as possible. However, to maintain a small oeffiient of moment, we need to keep the loation lose to the first quarter of the hord.[5.3] Therefore a ompromise was reahed in piking a amber at 40% hord length whih has a reasonably low pithing moment and provides high l values. When airfoil thikness inreases, the l values inrease. However, l values stop hanging signifiantly after 12% thikness to hord for NACA 4-digit airfoils. [4] Sine a greater thikness to hord ratio inreases drag and the wings of the UAV are not going to store fuel, an airfoil thiker than 12% hord is not pratial. Thus a thikness of 12% hord was hosen. The following four graphs were plotted from the output of XFOIL. Assumptions made inlude a Reynolds number, Re, of a onstant 2.3E5 and a Mah number, M, of onstant Plots of l versus alpha for five airfoils with inreasing amber are shown in Figure 5.1. This graph allows us to identify the lmax of the various airfoils whih is the maximum point for the five plots. April 5, 2007 Page 50 of 72

51 Cl vs. alpha for inreasing amber Cl NACA 0012 NACA 1412 NACA 2412 NACA 3412 NACA Alpha (degrees) Figure 5.1 C l versus Alpha for Inreasing Camber Figure 5.1 shows that an inreased angle of attak inreases the oeffiient of lift. The maximum point of eah line in the graph represents the lmax. The NACA 4412 has the highest lmax, 1.65 and l0, 0.5. Plots of m vs. alpha for five airfoils with inreasing amber an be seen in Figure 5.2. These plots show the flutuations in m as alpha is inreased. Cm vs. alpha for inreasing amber Cm Alpha (degrees) NACA 0012 NACA 1412 NACA 2412 NACA 3412 NACA 4412 Figure 5.2: C m versus Alpha for Inreasing Camber April 5, 2007 Page 51 of 72

52 As alpha hanges, the m flutuates as shown above. NACA 0012 has zero pithing moment when angle of attak, alpha, is zero beause it ats as a symmetri airfoil. However, as NACA 4412 produes the most lift, at l0, it also has the largest moment oeffiient. Plots of d vs. alpha for five airfoils with inreasing amber follow in Figure 5.3. The plots show how d inreases with an inreased alpha. Cd vs. alpha for inreasing amber Cd Alpha (degrees) NACA 0012 NACA 1412 NACA 2412 NACA 3412 NACA 4412 Figure 5.3 C d versus Alpha for Inreasing Camber The oeffiients of drag for the five airfoils are similar. This shows that an inrease in amber does not greatly affet the drag produed by eah airfoil. Plots of drag polar d vs. l for five airfoils follow in Figure 5.4. The optimal point for ruise would be the point with the highest l / d whih is the point furthest to the right for eah plot. April 5, 2007 Page 52 of 72

53 Drag Polar Cd Cl NACA 0012 NACA 1412 NACA 2412 NACA 3412 NACA 4412 Figure 5.4 Drag Polar Sine the drags for the five airfoils are similar, the airfoil with the highest lift to drag ratio, L/D, would be the NACA 4412 airfoil. Ultimately, the NACA 4412 airfoil was hosen for its high l0 and lmax, a low drag, a small enter of pressure movement aross large speed range, a reasonable pithing moment. 5.2 Wing Sweep Wing sweep serves the purpose of reduing transoni shok and supersoni flow. However, sine the UAV s maximum speed is below Mah 0.2, wing sweep would not be required as it would add to manufaturing ost. [5.4] However, if the aerodynami enter of the airraft needs to be moved bak far enough for balane, swept wings may be an option. 5.3 Taper Tapered wings are used to simulate an elliptial wing loading. An elliptial wing loading is desirable as it minimizes indued drag for a given span. Thus, a plot was generated on MATLAB from the following equations and assumptions to obtain a taper ratio for a minimum wing area. A taper ratio will save material and allow for a higher wing loading. From lifting line theory, [5.5]: April 5, 2007 Page 53 of 72

54 Assumptions inlude: elliptial wing loading, size of planform does not affet the weight of the airraft (weight is kept onstant at gross takeoff weight), effet of the anard is small, and steady level flight. Variables : C l =lift oeffiient Γ (y) =vortex distribution (y)=hord length along planform V=veloity b=span S=planform area L=W=603 lb (from onstrain analysis) l L' = 0.5ρV 2Γ( y) 2Γ 2 = = 1 (2y / b) 2 s ( y) V ( y) V ( y) y Γ( y) 2Γb 1 (2y / b) ( y) = s(1 ) α αlo = b 2V b πs (1 y / b) l 2 2Γ 1 (2y / b) 2Γs = = f ( y) 5.3 V 1 y / b V s For minimum wing area, f (y)=0,taper ratio =0.5 y =b/4 f(y=b/4)=(3/4) -.5 s 2 Γs Γs l V ( ) l = f y => = V s s 2 f ( y) 5.4 L Γmax b = 4/ π 2 ρv 5.5 Γmaxb = bmax Γs / s 5.6 S = 75b 2 68 ft 5.7 min. max = April 5, 2007 Page 54 of 72

55 Figure 5.5 Wing Area versus Taper Ratio When ompared to a wing without taper ratio, a savings of over 10% of material results by using a taper ratio of 0.5 (minimum point on graph). Therefore, the wing area is redued from 78 square feet to 69 square feet. From this wing area and an aspet ratio of 13 whih was obtained from the onstrain analysis, the resulting span (b) was 30 feet and the maximum hord length (s) was 3 feet. The maximum lift oeffiient (C lmax ) for the entire wing was also obtained as Twist To obtain an elliptial wing loading, wing twist must be added. [5.5] δ = k Where δ is twist and k is elliptial effiieny Sine the wing is tapered, the elliptial effiieny would be lose to 1, about [5.6] Thus we ll yield a wing twist of 0.04 radians or 2.23 degrees. Based on the taper ratio and wing twist, an elliptial wing loading an be obtained whih in turn redues indued drag. April 5, 2007 Page 55 of 72

56 6.0 Propulsion System The propulsion system for the Metro-Sout is a pusher piston propeller engine as mentioned in setion 3.3. To reiterate, it will be best loated at the bak of the airplane thus making it a pusher propeller type. This onfiguration is the best for the Metro-Sout beause the amera part of the payload is in the nose. The propeller blades need to be out of the line of site of the amera. Plaing the propulsion system in the front of the airraft would not only obstrut the view of the amera but it would also reate very turbulent flow around the amera reating extra vibrations and noise that will distort the piture. 6.1 Engine Seletion The initial horsepower requirement of the engine, 40 brake horsepower, was determined based on the weight, endurane, range, max speed, and power loading onstraints of the UAV. This value was then taken to the UAV database to find off-the-shelf engines that met the power requirement. Table 6.1 shows the list of engines that were initially onsidered. Engine Lightening Airraft AR 801 AR 801R Rotax 503 Rotax 582 Engines 604D4-F1 Power (bhp) RPM SFC Weight (lbs) Table 6.1: Engine List [6.1][6.3] The AR-801 is a Wankel-type rotary, single rotor engine with a apaity of 294, brake horsepower of 35-60bhp at 8000RPM, and a speifi fuel onsumption of 0.56 at max power. It was hosen for its size, weight, speifi fuel onsumption and brake horsepower. The AR-801 engine has dimension of 1foot x 1.06 feet x 0.82 feet. This engine is known to be a highly optimized, light-weight, single rotor, liquid ooled engine. It is designed suh that the mounting of alternators between 0.9 and 2.0 KW is feasible. It has been designed and developed speifially for UAVs requiring 35 to 60 bhp, with diret drive to propeller or vehile gearbox. Other engines built by the same ompany are urrently being used in other UAVs suh as the RQ-6 Outsider and RQ-7 Shadow-200. [6.2] Below is a list of the major advantages to the use of this partiular engine. Figure 6.1 is a piture of what the April 5, 2007 Page 56 of 72

57 AR-801 engine looks like with 4 blades. This is urrently the hosen engine for the Metro- Sout. Figure 6.1: AR-801 engine [6.1] Use of an AR-801 engine: Advantages: Team 4 has hosen the AR-801 UAV engine for the Metro-Sout for the following reasons: (1) High Power to Weight Ratio: A larger power to weight ratio allows for better speed ontrol and maneuverability of smaller airraft. It also helps to derease the overall weight of the airraft but still produing enough power to meet the power requirement. (2) Eonomi Fuel Consumption: An eonomi fuel onsumption allows for the airraft to fly farther per gallon of fuel used. This in turn an inrease the endurane time and range of the airraft, thus allowing for more ontinuous area overage. (3) Low Levels of Vibration: Vibration levels are extremely important when dealing with airraft design. The lower the levels of vibration, the less stress ats on the airraft. In the instane of the Metro-Sout, this means that although there is more stress on the tail setion, overall there is less stress on the airraft as a whole. (4) Low Cross Setional Area: The low ross setional area of the AR-801 engine helps to derease the amount of drag that is produed. On the Metro-Sout the engine is not streamlined into the fuselage, it is in fat, a separate entity that it attahed to the bak of the tail setion. In most ases the pusher prop engine reates a large amount of drag, but the lower the ross setional area of the engine, the less it stiks out around the fuselage, and less additional drag is reated. April 5, 2007 Page 57 of 72

58 (5) Long Life: Although the lifespan of the engine is not a requirement, it is a definite advantage for the engine to have a longer lifespan so that it doesn t have to be replaed often. Replaing engines is extremely expensive and time onsuming. This engine type supports a variable pith propeller. The variable pith makes it possible for the pilot to hange the blade angle of the propeller at will in order to obtain the best performane of the airraft engine. At take-off the propeller is set at the low blade angle so that the engine an attain the max allowable power and rpm. Shortly after take-off the angle is inreased slightly to prevent overspeeding of the engine and to obtain the best limb onditions of the engine rpm and airraft speed. When the airraft has reahed ruise or loiter altitude, the propeller an be adjusted to a omparatively high pith for low ruising rpm. This would allow for the Metro-Sout to be muh more adaptable to flight onditions in the instane of a high speed hase. [6.4] 6.2 Propeller Sizing Although the engine ame with a known size of propeller blades, they were too long for the urrent design of the Metro-Sout. The blades would have struk the ground on take-off and so in order fix this problem, the propeller blades were sized using the following method. Using the power required, many other parameters and speifiations of the engine were alulated. The advane ratio and the ativity fator are two very important parameters when understanding the blade design of the propulsion system. The advane ratio, found in equation 6.1, is just based on veloity, rotational speed and diameter of the blades. J = V nd (6.1) The advane ratio, muh like the wing angle of attak, is the related distane the airraft moves with one turn of the propeller. The advane ratio for the Metro-Sout is The ativity fator is a measure of the effet of blade width and width distribution on the propeller and is a measure of the propeller s ability to absorb power. Equation 6.2 is the equation of the ativity fator per blade. April 5, 2007 Page 58 of 72

59 AF perblade = 5 10 D R r R 3 dr (6.2) The average ativity fator for small, light airraft is approximately 100. The ativity fator for the blades on the Metro-Sout is 97. Equation 6.4 below shows how the thrust required was obtained. The oeffiient of thrust ( T ) was found using the propeller polar relation, shown in equation 6.3 and figure 6.2, between the power oeffiient and thrust oeffiient. T P = m + b (6.3) 2 2 J J Figure 6.2: Propeller Polar Plot for AR-801 Engine In equation 6.3, m and b are the slope and y-interept of the propeller polar plot. The power oeffiient ( P ) was found, seen in equation 6.4, sine the power required was already known. P = (6.4) n 3 D 5 P ρ T = (6.5) n 2 D 4 T ρ April 5, 2007 Page 59 of 72

60 In equations 6.4 and 6.5, ρ is the density of air at sea level, n is the rotation speed, and D is the propeller diameter. From the thrust equation, equation 6.5, the propeller effiieny was alulated to be 0.76 in equation 6.6. [2.1] TV η P = (6.6) 550* bhp 7.0 Longitudinal Stability Analysis 7.1 Stability Definition The basi onept of stability is simply that a stable airraft, when disturbed, tends to return, by itself, to its original state [2.1]. Stability is one of the important issues when building an airraft. There are some terms assoiated with stability whih are important to alulate and realulate for optimization. These inlude enter of gravity loation (.g), neutral point (n.p) and stati margin (SM). Early estimations of what these values should be an help in determining the urrent stability of the airraft. The methods of finding these variables are disussed here. Before disussing the methods, some symbols and aronyms are to be noted: M g - moment about.g M w - wing aerodynami pithing moment M - anard aerodynami pithing moment L - anard lift L w - wing lift x a w - aerodynami enter of the wing (with respet to wing L.E) x a - aerodynami enter of the anard (with respet to wing L.E) x g - enter of gravity loation α - airraft angle of attak i w - wing inidene angle i - anard inidene angle ε - average downwash angle indued by anard w - wing mean hord - anard mean hord q - dynami pressure S - wing area - airraft hord length (fuselage length) April 5, 2007 Page 60 of 72

61 Figure 7.1 represents the free body diagram used as its base model. Taking into aount all aspets of the fore system, suh as downwash and inidene angle, greatly ompliates deriving an initial formula for neutral point. Thus, designers made several key assumptions in determining the initial analysis: 1) Drag and thrust are negligible 2) Downwash and fuselage effets are negligible 3) α is relatively small (os(α ) 1) 4) q = q q w = ε 5) Change in downwash angle with α is negligible ( = 0 ) α 6) α α = α w = Figure 7.1Free Body Diagram [7.1] 7.2 Neutral Point Calulation The team set out to evaluate stability based on the loation of the neutral point, or the point on the airraft about whih the net moment does not hange with angle of attak [7.2]. This method essentially finds the point at whih the airraft enter of gravity,.g., rests in relation to the airraft aerodynami enter. Figure 7.2 diagrams the fores on the airraft and the loations of these fores with the designers key assumptions in mind. Note that the referene loation for this analysis rests at the leading edge of the wing. April 5, 2007 Page 61 of 72

62 Figure 7.2: Fores and Referene Points for Neutral Point Based on figure above, the team derived equation 7.1 for moment about the airraft.g.. M g w ( x a x g ) L w + M + ( x g x a ) L = M 7.1 Originally, the equation looked like equation 7.2. M g w w ( x x ) L ( α + i ε ) + M + ( x x ) L os( i ) = M os α 7.2 aw g w w g a + In equation 7.2, the group assumed α, i w and ε were small ( 0 ) and thus the osine term equals one, leading to Equation 7.1. Next, designers altered equation 7.1 to reate Equation 7.3 by dividing by airraft mean hord, dynami pressure, and wing area. This led to the non-dimensional oeffiient form found in equation 7.4, where M M M m = and qs L L =. qs ( x x ) ( x x ) L M S g w w aw g w g a = qs w qs qs S qs S qs ( x x ) ( x x ) w a S S w g g a m, g = m L + m + L 7.4 w w S S L S Analysis required multiplying eah moment term in Equation 7.3 by the respetive hord length of its omponent ( w / w for the wing and / for the anard) to obtain the proper April 5, 2007 Page 62 of 72

63 April 5, 2007 Page 63 of 72 moment oeffiients for those omponents. Equation 7.5 represents Equation 7.4 after taking the derivative of eah term with respet to its orresponding omponent s angle of attak. ( ) ( ) L a g m w w L g a w w w m m g S S x x S S x x w α α α α α α α α α α,,,,, + + = 7.5 Again, the following terminology applies: w α wing angle of attak α anard angle of attak. However, α α α = = w is one of the assumptions. Taking a derivative and setting equation 7.5 to be equal to zero yields equation 7.6. ( ) ( ) S S x x S S x x L a g m w L g a w w m m g w α α α α α,,,,,, = = 7.6 The goal here is to find x g, hene rearranging equation 7.6 yields equation = S S S S S S x x x L w L m L a w w m w L a g w α α α α α α,,,,,, 7.7 Finally, the team divided eah term on the right hand side of equation 7.7 by α w L,, and the final equation beomes equation = S S S S S S x x x w L L w L m w L L a w w L w m a g w α α α α α α α α,,,,,,,, As mentioned before, this x g is atually the airraft neutral point (n.p) ( np g x x = ) with respet to wing leading edge as the referene point. This implies that stati margin is zero when np g x x = (here, g x is the atual.g. loation). Theoretially,.g. an be forward

64 (ahead) or aft of the neutral point. To have a positive stati margin whih makes the airraft stable, the.g must be ahead of n.p based on the stati margin formula in Equation 7.9. xg SM = x g x np = x g 7.9 The variable x g is the atual.g. loation. By saying atual, this.g. is derived from the following statistial group weight method. The next step in determining the airraft s stability involves determining the loation of its enter of gravity. The team used the statistial group weighted method summarized by equation 7.10 to aomplish this task. _ g x n i= 1 = n W i= 1 i W x i _ i 7.10 Equation 7.10 sums the individual produts of the weight of eah main airraft omponent (wings, anards, fuselage, et.) at its respetive enter of mass and the omponent s distane from the leading edge of the airraft. It then divides the result by the sum of the individual weights of the omponents. Figure 7.3 presents a visual model of this method. In a dynamial analysis, this method essentially treats eah omponent as a partile mass loated some distane from a referene point, whih in this ase, is the leading edge of the airraft. Payload Canard Vertial Engine Tai Fuel Wing Fuselage Figure 7.3: Geometry for Finding Airraft Center of Gravity April 5, 2007 Page 64 of 72

65 Center of gravity depends upon fixed weights, suh as those for the struture and payload, along with variable weights, suh as that of fuel. Therefore, as the plane uses fuel during the flight, the enter of gravity will shift, thus hanging the stati margin. For this reason, the team will plae the tank ontaining a majority of the fuel as lose to the enter of gravity as possible to keep the enter of gravity and, therefore, the stati margin, from moving out of tolerane ranges during flight. 8.0 Summary Thus far, Team 4 has determined to provide a primary ustomer base omprised of polie and news organizations with the Metro-Sout, an unmanned aerial vehile apable of performing those tasks for whih those ustomers urrently use onventional heliopters. This raft will perform both autonomously and with a remote pilot, depending on the mission suh as safe operation at ft above ground level, a overage radius of 200mi, an endurane of at least five hours, and a payload weight of between sixty and seventy pounds. To perform suh objetives, the team has determined key design attributes as outlined. The group aims to sell the Metro-Sout to target ustomers at a lower aquisition and operating ost than urrent heliopters to be ompetitive within the market. After onsidering several possible designs for the airraft, the team deided on a anard onfiguration with a front mounted amera and pusher piston propeller that would best aomplish the neessary missions. Airfoil seletions as well as wing shape are urrently in plae to provide adequate lift, and the airrafts longitudinal stability established by onfirming the plaements of both the enter of gravity and the neutral point. The next step forward in the design proess involves several elements. First, the team must deide on an aspet ratio, whih optimizes both aerodynami need and strutural limitations. There are also several aspets of the airraft that will need additional optimization iterations in April 5, 2007 Page 65 of 72

66 order to reate a better performane the required mission. The airraft will need to be modeled, tested, re-evaluated, and eventually produed and distributed. April 5, 2007 Page 66 of 72

67 R.0 Referenes [1.1] "Bell Heliopter Commerial." Bell Heliopter Textron In. 28 Jan [1.2] AOPA, AOPA s 2006 Aviation Fat Card, download.aopa.org/epilot/2006/fatard.pdf [retrieved 18 February 2007]. [1.3] The Heliopter Market Newsletter Piston, Heliopter International Assoiation, published Nov. 28 th, 2005, [retrieved February 18th, 2007] [1.4] The Heliopter Market Newsletter Turbine, Heliopter International Assoiation, published Ot. 25th, 2005, [retrieved February 18th, 2007] [1.5] Clark, Larry K., The Life of a Pilot Flying ENG (eletroni news gathering), justheliopters.om (online), 2/21/2004, el=artiles, [retrieved February 15 th, 2007] [1.6] 2000 Airraft Industry Studies, The Industrial College of Armed Fores, [retrieved Feb. 15th, 2007] [2.1] Raymer, Daniel. P., Airraft Design: A Coneptual Approah (4 th Ed.), Blaksburg, VA: AIAA In, (2006). [2.2] Kroo, Ilan., Shevell, Rihard., Chapter 12.1, Airraft Design: Synthesis and Analysis, Retrieved April 1 st, 2006, from Desktop Aeronautis In., (2006). [2.3] Lennon, Andy., Canard: A Revolution in Flight (1 st Ed.), Hummelstown, PA: Aviation Publishers, (1984). [3.1]Clausing, Don. Total Quality Development. New York: ASME P, [3.2] Mobile polie radar gun simion speifiations page [Retrieved 10-Feb-07]. [3.3] Pixel 275 III polyteh speifiations page [Retrieved 10-Feb-07]. April 5, 2007 Page 67 of 72

68 [3.4] ThermaCAM SC3000 Flir Systems therma am speifiation page [Retrieved 10-Feb-07]. [3.5] Hal, Sony Dvam DSR-PD150 DV Speifiations produt wiki database 0_dv_speifiations.html [Retrieved 10-Feb-07]. [3.6] 3400 Full featured UAV autopilot 3400 Auto pilot speifiations [Retrieved 10-Feb-07]. [3.7] Speifiation PowerShot S3 IS S3 IS speifiations =144&modelid=13077 [Retrieved 10-Feb-07]. [3.8] Speifiations EF mm f/ II USM EF f/ II USM speifiations modelid=9435 [Retrieved 10-Feb-07]. [3.9] C I N E F L E X V 1 4 M A G N U M - M U LT I S E N S O R Cineflex 14 M Speifiations [Retrieved 10-Feb-07]. [5.1]Husa, B Airfoil Seletion [Retrieved April 2nd 2007] [5.2] NACA Airfoil Series [Retrieved April 2nd 2007] [5.3]Heperlle, M Loation of Camber and Moment Coeffiient [Retrieved April 2nd 2007] [5.4]Abbott, I.H., and von Doenhoff A. E., Theory of Wing Setions: Inluding a Summary of Airfoil Data, Dover, pp 288. [5.5]Lyrintzis, A.S., AAE 514 Intermediate Aerodynamis Spring 2006, Copymat, West Lafayette, 2006,pp [5.6] Indued drag, [Retrieved April 2nd 2007] [6.1] "AR801-50bhp." UAV Engines Ltd UAV Engines Ltd. 27 Feb April 5, 2007 Page 68 of 72

69 [6.2] "Shadow-200." Unmanned Aerial System AAI Corp. 25 Mar [6.3] "Rotax 503UL DCDI." Kodiak Researh Ltd. 6 Mar [6.4] Bent, Ralph D., and James L. MKinley. Airraft Powerplants. 5th ed. NY, NY: MGraw- Hill Book Co., [7.1] Roskam, Jan. Airplane Flight Dynamis and Automati Flight Controls. 4th ed. Lawrene, KS: DARorporation, [7.2] Steve, Brandt. Introdution to Aeronautis: a Design Perspetive. Reston, VA: Amerian Institute of Aeronautis, April 5, 2007 Page 69 of 72

70 A.0 APPENDIX A.1 Projet Timeline Desription Team 4 eleted to develop a projet timeline to establish a baseline measure of progress through the ourse of the semester. Team 4 has speifially targeting a number of phases in the design for overlap to allow the team greater freedom to make design hanges and foster greater ustomer partiipation in formulating design requirements. For instane, the projet timeline shows that Customer Attribute Identifiation phase goes hand in hand with the Initial Coneptual Design phase until the date of the Systems Requirements Review whereat all the ustomer attributes need to be finalized. The same is true for ertain aspets of the Initial Coneptual Design and the Design Analysis and Tweaking phases. The premise behind the layout of the timeline is to establish onstraints and deadlines that keep Team 4 moving forward in the design proess while giving it the freedom to make hanges as deemed neessary to keep the projet ompetitive. The five main stages in Team 4 s timeline and their urrent progress are (1.) Establish Customer and Produt: Phase Complete (2.) Customer Attribute Identifiation: Phase Complete (3.) Initial Coneptual Design Phase Ative (4.) Design Analysis/Tweaking Phase Ative (5.) Design Finalization Planned/Not Ative April 5, 2007 Page 70 of 72

71 Figure A.1: Gantt hart Team 4 s Projet Timeline April 5, 2007 Page 71 of 72