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Çift turboprop motorlu ve sivil amaçlı bir keşif/devriye uçağının roskam yöntemi ile ön dizaynı

Başlık çevirisi mevcut değil.

  1. Tez No: 46356
  2. Yazar: SELÇUK ESEN
  3. Danışmanlar: PROF.DR. A. NURİ YÜKSEL
  4. Tez Türü: Yüksek Lisans
  5. Konular: Uçak Mühendisliği, Aircraft Engineering
  6. Anahtar Kelimeler: Belirtilmemiş.
  7. Yıl: 1995
  8. Dil: Türkçe
  9. Üniversite: İstanbul Teknik Üniversitesi
  10. Enstitü: Fen Bilimleri Enstitüsü
  11. Ana Bilim Dalı: Belirtilmemiş.
  12. Bilim Dalı: Belirtilmemiş.
  13. Sayfa Sayısı: 123

Özet

ÖZET Bu yüksek lisans tez çalışmasında, GAP Projesi çerçevesinde Güneydoğu Anadolu Bölgesi'nde sivil amaçlı keşif / devriye görevi yapacak, mürettebat dahil beş kişilik, çift turboprop motorlu, küçük yapılı bir uçağın Roskam Yöntemi 'ne göre ön boyutlandırması ve Roskam 'm Sınıf- 1 Yöntemi ile ön konfigürasyon dizaynı yapılmıştır. Sözkonusu uçak, görev ve performans özellikleri açısından bir sivil keşif / devriye uçağı, genel görünüş ve ağırlık özellikleri açısından ise bir bölgesel (regional) turboprop 'tur. Görev bölgesi ve yerine getireceği görevin niteliği gözönüne alınarak“GAP Devriyesi”adı verilen ve öncelikle görev spesifikasyonlan belirlenen bu uçağın, ön boyutlandırma hesapları sonucunda, performans ve boyut karakteristikleri (maksimum kalkış ağırlığı, maksimum kalkış motor gücü, kanat açıklık oranı, boş ağırlık, görev yakıt ağırlığı, paralı ağırlık, maksimum taşıma katsayıları ve kanat alam) belirlenmiş, Sınıf- 1 yöntemine göre ön konfigürasyon dizaynı sonucunda ise kokpit ve gövdenin genel görünüşü, güç grubu ve uçak üzerindeki yerleşimi, kanat planformu, yanal kontrol ve aşırı taşıma yüzeylerinin boyutları ve yerleşimi, kuyruk ile kuyruk kumanda yüzeylerinin boyutları ve yerleşimi ortaya çıkmıştır. Son olarak, elde edilen tüm veriler ve hesaplanan tüm değerlere göre sözkonusu uçağın üç yönden (önden, üstten ve soldan) görünüşü çizilmiştir. Sınıf- 1 yöntemine göre ön konfigürasyon dizaynının son kısmım oluşturan iniş takımları ayrıntılı dizayn ve yerleşimi, ayrıntılı ağnlık-balans, stabihte-kontrol ve sürükleme polar analizleri ise bu tezin kapsamına dahil edilmemiştir. Ön konfigürasyon dizaynının daha ayrıntılı aşamalarım içeren Sınıf-2 yöntemi, bireysel çalışmayla gerçekleştirilmesi mümkün olmayacak kadar ağır ve kapsamlıdır. Bu nedenle Sınıf-2 yöntemi GAP Devriyesi 'nin ön dizaynında yalnızca gerektiği yerlerde danışılan bir yardıma unsur olarak kullanılmıştır. Bu yüksek lisans tezinde İngiliz (Anglosakson) birim sistemi kullanılmıştır. Bunun iki temel nedeni vardır: 1-) Dizayn için seçilen Roskam Yöntemi 'nde tüm yaklaşım ve hesaplamalar İngiliz birim sistemiyle yapılmıştır. Mümkün olduğu kadar yuvarlatılmış rakamlı sonuçlar elde edebilmek ve bu sonuçlan Roskam'ın örnek sonuçlarıyla karşılaştırabilmek amacıyla sözkonusu birim sisteminin kullanılması zorunlu olmuştur. 2-) İngiliz birim sistemi, dünya sivil havacılığında çok yaygın olarak kullanılmaktadır. Özellikle İngiliz ve Amerikan kaynaklı dünyanın en büyük sivil uçak üreticileri (British Aerospace, Mc Donnell Douglas, Boeing vs.), yayınlarının hemen hemen hepsinde bu sistemi temel almaktadırlar. İngiliz ve SI birim sisternlerinin birbirlerine dönüşümlerini gösteren tablo Ek A 'da sunulmuştur. XX

Özet (Çeviri)

PRELIMINARY DESIGN OF A TWIN TURBOPROPELLER ENGINE DRIVEN CIVIL PATROLLING AIRPLANE VIA ROSKAM METHOD SUMMARY In this study preliminary sizing and preliminary configuration design of a twin turbopropeller engine driven civil patrolling airplane have been performed using Roskam 's Method. The subject airplane, which is called“GAP Devriyesi”regarding its mission region (because it is being planned to be charged in Southeast Anatolia for Southeast Anatolian Project), is a civil patrolling airplane from the point of view of its mission and performance specifications whereas it can be treated as a regional turbopropeller engine driven airplane from the point of view of its general configuration outline and weight data. For the sake of accurate and logical improvement of estimations and calculations during all steps of the design, a sample airplane has been selected: Beechcraft Super King Air B200T. This airplane is a proper sample to be referred since it can be classified in the same manner like GAP Devriyesi: A twin turbopropeller engine driven as well as a civil patrolling airplane. Weigt, performance and configuration data of B200T have frequently been used as reference during this study and they have been compared to those of GAP Devriyesi step by step. Roskam 's Method of preliminary sizing begins with determination of mission specifications for the airplane to be designed. GAP Devriyesi has the following mission specifications: 1- Payload (including crew): 1700 lb 2- Range (with 15 % reserve fuel and maksimum payload): 2000 sm. 3- Flight Altitude (from sealevel): a) patrolling altitude: 5000 ft, b) maximum cruising altitude (with all engines operative): 35000 ft (= service ceiling) 4- Speed: a) patrolling speed: 260 mph, b) maximum cruising speed at 14000 ft: 305 mph 5- Climb Rate (at sealevel and with all engines operative): a) 2500 ft/min, or b) 2000 ft/min 6- Patrolling Time: 8 hours 7- Take-off and Landing: a) altitude: 2000 ft (from sea level - Diyarbakir Airport), b) landing performance at WL= 0.90 Wxo, c) take-off ground run: 1800 ft, d) landing ground run: 1450 ft 8- Wing Aspect Ratio: 10 9- Powerplant: 2 x turboprop 10-Pressurization: 8000 ft cabin altitude at 35000 ft 11-Certification Base: FAR 23 The relationship between maximum take-off weight (WTO), payload (WPL), mission fuel weight (Wf) and operational empty weight (Wqe) is given as: XXIWTo = WpL+WF+WoE (1) Initial estimate of WTO for GAP Devriyesi is: Wto = 14000 lb Mission fuel weight has been calculated via Roskam 's Fuel Fraction Method as: WF = 0.300 WTO (2) Defining a tentative value for empty weight (Who*), symbolizing weight of trapped fuel - oil as Wtf0 and regarding a linear relationship between logio Wto and logio WE, following five equations have been obtained: Wtf0 = 0.005 WTO (3) WoHeat = WEtalt+Wtfo (4) WF = 0.300 WTO (2) WTO = WPL + WF + WoEtmt (5) WE = inv. logi0{(log10WTo-A)/B} (6) where A and B are regression line constants. Using these equations with an iterative approach, final values for characteristic weights of GAP Devriyesi have been revealed: WTO = 14358 lb WE = 82791b WF = 43071b In order to determine the sensitivity of Wto to the changes in a) Payload,WpL b) Empty weight, WE c) Range, R d) Patrolling time, tpt e) Patrolling speed, Vpt f) Specific fuel consumption, Cp g) Propeller efficiency, tiP h) Lift-to-drag ratio, L/D sensitivity studies have been performed and following relations have been obtained: 5Wto/9Wpl= 10.28 lb/lb öWto/3We= 1.67 lb/lb XXll5WTO/3R = 14.49 lb/sm oWTO/5tpt = 3766 lb/h öWTO/öVpt = 1161b/mph aWTO/3cp = 50218 lb/lb/hp/h 5WTO/^nP = -39131 lb flWTO/5(L/D) = -16741b In the next step, calculations which meet: a) Stall speed requirements b) FAR 23 take-off distance (sro) requirements c) FAR 23 landing distance (sL) requirements d) FAR 23 climb requirements, with drag polars: CD = 0.025 1 + 0.0374 (CL)2 (clean config.) (7) CD = 0.0401 + 0.0413 (Cjf (take-off config., landing gear retracted) (8) CD = 0.065 1 + 0.0413 (CL)2 (take-off config., landing gear extended) (9) CD = 0.0901 + 0.0442 (CjJ2 (landing config., landing gear retracted) (10) CD = 0.1151 + 0.0442 (Cl)2 (landing config., landing gear extended) (11) e) FAR 23 rate of cimb (RC) requirements f) FAR 23 climb gradient (CGR) requirements g) Time-to-climb requirements h) Ceiling requirements i) Cruise speed requirements have been fulfilled and it has been shown (via graphics) how take-off wing loading {(W/S)ro} changes with take-off power loading {(W/P)ro} for each of these cases. Consequently, all critical and effective results have been combined on a single graphic. Regarding the maximum allowable (W/S)ro and (W/P)ro values on this graphic, preliminary sizing of GAP Devriyesi has been completed and following results have been obtained: Take-off wing loading, (W/S)ro = 47.9 lb/ft2 Take-off power loading, (W/P)ro = 8.2 lb/hp Maximum take-off power, PTO = 1750 hp Wing area, S = 300 ft2 Maximum clean lift coefficient, Cl^ =1.5 Maximum landing lift coefficient, (Ci^^Jl - 2.5 Maximum take-off lift coefficient, (CLmax)!» =1.8 which are in good agreement with the data of sample airplane Beechcraft B200T. Preliminary configuration design of GAP Devriyesi has been performed via Roskam 's Class-1 Method. Roskam 's Class-2 Method, which involves more XXllldetailed investigations and requires significiant expenditure of engineering manhours, has been referred to and used only where necessary. The objective of Class- 1 method is to decide on the feasibility of a given configuration with a minimum engineering work. This method has limited accuracy. The objective of Class-2 method is to arrive at a reasonably detailed layout of a given configuration so that its mission capabilities can be compared to those of other competing concepts with confidence. This method has fairly good accuracy. Configuration design via Class- 1 method amounts to making the following decisions: 1- Selection of the overall configuration 2- Selection of the fuselage layout 3- Selection of propulsion system type 4- Selection of the number of propellers and determination of the propeller diameter 5- Integration of the propulsion system 6- Selection of planform design parameters for the wing and for the empennage (tails) 7- Selection of type, size and disposition of high lift devices (flaps and slats) 8- Selection of landing gear type and disposition 9- Selection of major systems to be employed by the airplane and performing weight and balance analysis 10-Performing stability and control analysis 1 1 -Determination of the drag polars In this study the first seven steps outlined above have been worked out and consequently threeviews of GAP Devriyesi have been obtained. The last four steps require so many engineering manhours and detailed investigations (although they are included in Class- 1 method) that they can separately be the subject of another master of science study. Ignoring these steps does not affect the reliability of the preliminary configuration design which has been performed by this study. The first seven steps are sufficient to reveal the general outline of the configuration design for GAP Devriyesi GAP Devriyesi is a land based airplane and is of a conventional configuration (ie. tail aft). It has a low wing layout like most regional turbopropeller driven airplanes have and its empennage configuration is T-taiL Its powerplant is of tractor type (propeller is forward of the airplane 's center of gravity) like all regional turbopropeller driven and most patrolling airplanes are. Its landing gear is of hydraulically retractable (tricycle) type with twin wheels on each main unit and single wheel on steerable nose unit. Nose landing gear is retractable rearward and main landing gear forward. GAP Devriyesi has been designed to carry five people (two crew: one pilot and one copilot, three observers) and 825 lb of total luggage plus portable patrolling and surveying equipment. One pilot and one copilot seats are located side to side in the cockpit. A one abreast of three single observer seats (on the left side of the cabin section) layout has been selected. XXIVA modular lavatory is installed on the right side of the rear vestibule. A narrow wardrobe / dry galley is installed on the right side of the forward cabin section. A pressurized bulk cargo compartment with 50 ft3 capacity is located in the tailcone. An outward opening emergency exit is installed on center outboard side overwing. Wall mounted book shelves are installed on the right side of the cabin section (one between wardrobe and emergency exit, one between emergency exit and lavatory). Internal cabin dimensions are as follows: Length: 17.6 ft Max width: 4.5 ft Max height: 4.8 ft External fuselage dimensions are as follows: Overall length (from nose to the end of tailcone), If = 39.5 ft Fuselage diameter, df = 6.4 ft Fuselage fineness ratio, Vdf =6. 17 Tailcone fineness ratio, lfC/df = 2.86 As a mission specification a twin turboprop engine configuration has been selected. Regarding the maximum power level per engine as Pmax = PTo/2 (12) three bladed Pratt & Whitney PT6A-42 turboprop engine which supplies a take-off power of 850 shp at 2000 propeller rpm to 41 °C has been selected for GAP Devriyesi Propeller diameter (Dp) has been calculated as 8.3 ft. Engines are buried in nacelles which are located in the left and the right wings. Structural wing configuration is of cantilever type and quarter chord sweep angle of the wing (Ac/4) is 0°. Wing thickness ratios are: At the wing centerline, (t/c)r = 0. 18 At the wing tip, (t/cX = 0. 12 Regarding these data and lift coefficient characteristics, NACA 23018 and NACA 23012 airfoils have been selected respectively for wing root (centerline) and wing tip. Airfoils are thinned from root to tip in a linear manner. Wing taper ratio (A^) has been considered to be 0.43. Using the following relationship b = (AS)ia (13) wing span (b) has been calculated: b = 54.8 ft Under these circumstances wing chords at root (c,) and at tip (ct) have been determined as XXVCr=6.72ft q = 2.89 ft It has also been shown that NACA 23018 and NACA 23012 airfoils meet minimum“section maximum lift coefficient”requirement. The requirement (demand) is (Clmax)r + (C]max)t = 3.38 and the supply is (Clmax)NACA 23018 + (Ctoiax)NACA 23012 = 3-22 which is acceptable within a deviation of 4.7% (Roskam allows maximum 5% deviation). GAP Devriyesi has been designed to have two single slotted flaps and one retractable slat on each wing and it has been proved that these high lift devices are able to meet the requirements of section lift coefficient increments during take-off and landing. Other characteristics of the high lift devices have been determined to be: a) Flap angle, Sf = 15 ° (during take-off), o> = 40 ° (during landing) b) Flap chord ratio, Cf/c = 0.20 c) Slat chord ratio, cs/c = 0. 195 d) Ratio for flapped wing area, Swf/S = 0.68 e) Flap span ratio, bf/b = 0.71 f) Slat span ratio, bs/b = 0.65 Lateral control surfaces (ailerons) of GAP Devriyesi are located between 80 % and 100 % of half wing span on each wing and aileron chord ratio has been assumed to be 0.21. Mean geometric chord of the wing (c ) has been calculated as 5.85 ft. The required fuel volume (vwf) which can be carried in the wing has been calculated via following equation: vWF = 0.54(S2/b)(t/c)r{(l + Xw(Tw)lfl + (Xw)2Tw)/(l + ^)2} (14) where Tw = (t/c)t/(t/c)r (15) vwf has been found to be 1 15. 1 ft3 and it has been shown that this volume is sufficient for all mission fuel (Wp = 4307 lb) to be carried in the wing. No wing tip tanks are necessary. Wing dihedral angle (Tw) has been selected to be 6°. Wing twist angle (st) is defined as St = (İw)t-(İw)r (16) XXVIwhere (vX is wing incidence angle at the wing tip and (iw)r is wing incidence angle at the wing centerline. For GAP Devriyesi CwX = -ı.ı° Öw)r = 3.80 Therefore; £t = -4.9° and is called“wash-out”(negative twist) which tends to supress tip stall. Assumptions and calculations for empennage (vertical and horizontal tail surfaces) sizing and disposition via Roskam 's Class-1 Method can be summarized for GAP Devriyesi by the following list: a) Moment arm of the horizontal tail, Xh = 24.6 ft b) Moment arm of the vertical tail, xv = 20.5 ft c) Horizontal tail area, Sh = 64.9 ft2 d) Vertical tail area, Sv = 52. 1 ft2 _ e) Horizontal tail volume coefficient, Vh( = XhSh/(Sc )) = 0.91 f) Vertical tail volume coefficient, vv( = xvSv/(Sb)) = 0.065 g) Elevator - horizontal tail area ratio, SJ\ = 0.28 h) Rudder - vertical tail area ratio, Sr/Sv = 0.29 i) Elevator root chord, (Cr)e = 0.42 Ch j) Elevator tip chord, (ct)e = 0.42 Ch k) Rudder root chord, (CrX = 0.47 Cv 1) Rudder tip chord, (ctX = 0.41 Cv Other geometric parameters are: For horizontal tail: a) Aspect ratio: 5.0 b) Leading edge sweep angle: 25 ° c) Taper ratio: 0.40 d) Thickness ratio: 0.12 e) Airfoiltype:NACA0012 f) Dihedral angle: 0 ° g) Incidence angle: fixed h) Root chord, (0,^ = 5. 14 ft i) Tip chord, (ctK = 2.06 ft_ j) Mean geometric chord, c h = 3.72 ft k) Span, bh = 18.01 ft For vertical tail: a) Aspect ratio: 1.0 b) Leading edge sweep angle: 40 ° c) Taper ratio: 0.65 d) Thickness ratio: 0.15 e) Airfoiltype:NACA0015 f) Dihedral angle: not applicable g) Incidence angle: 0 ° h) Root chord, (Cr)v = 8.75 ft xxvni) Tip chord, (ct)v = 5.69 ft_ j) Mean geometric chord, c v = 7. 17 ft k) Span, bv = 7.22 ft XXVM

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