Aircraft Design: Turkish Airports Specifications - Assignment Example

AIRCRAFT DESIGN NAME COURSE COURSE CODE INSTITUTION INSTRUCTOR DATE OF SUBMISSION TABLE OF CONTENTS INTRODUCTION SUMMARY OF TURKISH AIRPORT’S SPECIFICATIONS AIRCRAFT DATABASE AIRWORTHINESS REGULATIONS AND DESIGN WEIGHTS THRUST TO WEIGHT AND WING LOADING WING DESIGN HIGH LIFT DEVICES REFERENCES INTRODUCTION: Aircraft design is a very crucial activity in the aerospace industry as every aircraft has its own distinctive characteristics depending on its purpose and requirements. Aircraft design is a process undertaken by several professional design engineers prior to the final manufacture of the aircraft. This process aims at providing the engineers with an overview of the aircraft’s specifications and requirements to avoid increased manufacturing costs due to poor aircraft structures. Aircraft design is quite a complex process requiring an individual’s creative and analytical capabilities during the various design phases. Each phase of the aircraft design process has its own specified design methods and data requirements (Fielding 1999, p2). The first step involves assessing the requirements of the aircraft and its purpose. Depending on these factors, the aircraft design should seek to meet the standards set up by the respective flight authorities that aim at ensuring air safety. To ensure that all aircraft’s requirements are taken into consideration, three major aircraft design phases have to be achieved. Fielding identifies them as the conceptual, preliminary and detail design phases that are undertaken systematically (1999). For this module, the class was divided into several groups that were assigned various aircraft design projects that required information gathering. Upon the project’s completion, the group members was required to write individual personal reports and thereafter join the other group members in preparing the group report. This report discusses the personal progress carried out on the aircraft design project developed by the author. The author is (MEMBER 1) is a member of a group consisting of five members namely: (MEMBER 1) (MEMBER 2) (MEMBER 3) (MEMBER 4) (MEMBER 5) The aircraft chosen by our group is a 5-seater aircraft including the pilot. The range of the aircraft is 1300 nautical miles with a cruise altitude of 28,000 ft flying from Portsmouth Airport to Turkey. Each group member was assigned a specific task and time frame to allow other group members to continue with their respective sections of the group assignment. Initially, my role within the group involved researching on aircrafts with similar specifications. Later on I was handed the responsibility for T/W and W/S as well as weight estimates using an Excel sheet provided by the lecturer. SUMMARY OF TURKISH AIRPORT SPECIFICATIONS: According to my assigned areas of the project, I researched on the airport runways for all the international airports in Turkey. Eventually, my results indicated that the shortest runway was 1825 meters and the longest runway was 3750 meters. All the international airport runways in Turkey are listed in the table below with the main parameters of the runways. Name of Airport City Length of main runway in feet Length of Main Runway in Meters 1 Adana Sakirpasa Airport Adana 9022 2750 2 Esenboga International Airport Ankara 12303 3750 3 Antalya Gazipasa Airport Antalya 5988 1825 4 Dalaman Airport Dalaman 9842 3000 5 Ataturk International Airport ISTANBUL 9843 3000 6 Izmir Adnan Menderes Airport IZMIR 10630 3240 7 Trabzon Airport Trabzon 8661 2640 From this table, it was concluded that the maximum field length for a landing in all Turkish international airports has to be less than the shortest runway. The shortest runway is about 1825 meters, therefore it is safe to estimate the landing distance at 690 meters. AIRCRAFT DATABASE: The table below indicates similar aircrafts researched prior to the design process after having understood clearly the objectives of the group’s task. Aircrafts included in the table have been assessed from various established and promising manufacturing companies. AIRCRAFT Diamond D-Jet Epic Victory Piper Jet Cruiser Jet MANUFACTURER Diamond Aircraft Epic Aircraft Piper Aircraft Maverick Jets ENGINES 1× Williams FJ33-4 1× Williams FJ33-4 1× FJ44-3AP 2× Williams FJ33-4 SEATERS 5 4 to 6 7 4 MAX. CRUISE 315 knots 320 knots 360 knots 472 knots MAX. RANGE 1,350 NM 1,200 NM 1,300 NM 1,450 NM AIRWORTHINESS REGULATIONS & DESIGN: According to Torenbeek (1976), airworthiness regulatory requirements are imposed on airplane manufacturers with the aim of providing aircraft standards that seek to enhance air safety. This emanates from conclusions drawn from several air accident investigations that identify the basic aircraft structures as the major source of flight malfunctions and accidents. These standards in turn impact on the design of the aircraft’s structure, systems performance, installations and its flying capabilities. In designing an aircraft, engineering designers have to choose and implement the respective airworthiness code. Standards pertaining to aircraft design vary across nations as well as airplane categories and they should be observed strictly. Due to the regulatory variations, nations have established distinctive aviation authorities that ensure the aviation standards in the nation are observed. Other countries have similar rules with some troublesome differences that eventually result to confusion and incurring extra costs when different rules are enforced. The figure below indicates the classification of aircraft categories in the USA and UK. Figure 1. Classification of aircraft categories in US and UK Air Worthiness Requirements (Torenbeek, 1976) Aircraft designs have to choose the group, category or class that the aircraft belongs prior to the commencement of the design process. In the table above, the upper section shows the division of aircrafts into groups while the lower section indicates the appropriate UK and US requirements. In the UK, civil aircrafts are classified as light while in the US their classification is dependent on the maximum takeoff weight. When the maximum takeoff weight is less than 12500 pounds, the aircraft is classified as small. FAR regulations and airworthiness for design are sub-divided into the following sub parts. The designer has to ensure that the requirements for all these sub parts are satisfied individually. Torenbeek highlights several design considerations such as takeoff weight, bird proof windshields, runway landing performance, passenger capacity and weight allowance for each passenger, a minimum of two pilots if the number of passengers exceeds 9 among other design considerations (1976). Figure 2. Airworthiness and Design differences for small and transport category (Torenbeek, 1976) It is expected that the design would go through a number of configuration and iterations to meet the design requirements of payload, number of passengers, range, takeoff weight and other requirements. WEIGHTS: Being a group assignment, the weights estimates were calculated individually by the group members to provide an average of best results for the aircraft design project. Afterwards, the best values were decided upon comparison of existing aircrafts following research work conducted on them. My work in weight estimates and spread sheet which was given from the lecturer is shown here: Range R 1300 nm 2407600 m Specific fuel Consumption of the Engine 0.0001944 (N/s)/N Max. Mach Number M 0.7567 Aspect Ratio AR 6 Passenger Include Pilot PAX 5 Length of Fuselage L 12.5 m Breadth of Fuselage 1.5 m Height of Fuselage 1.5 m Lift to Drag (L/D) 12.12497 Cruise altitude - 8500 m Pressure at Cruise Altitude P 33.099 Relative temperature 0.80826 Fig.3: Weight Estimates After calculating and estimating values, I prepared a weights estimate table to calculate several important values as shown below: Aircraft Weight Estimates: Method 1 Torenbeek Method Cd0 = 0.02 R = 2407600 sfc = 0.0001944 theta = 0.80826 M = 0.7567 AR = 6 P = 33.099 L = 12.5 bf = 1.5 hf = 1.5 Wto = 44.03262 kN kcf = 4 Pax = 5 L/D = 12.12497 T/W = 0.255083 We = 2.246397 kN X = 1.936543 b = 0.149 c = -0.00782 ff = 0.316421 Ws = 9.14183159 Wto = 44.03262 kN Woe/Wto = 0.572184 Wp/Wto = 0.111395 Fig 4 According to the project instructions, the maximum takeoff weight must not exceed 45kn. The maximum takeoff weight was calculated with the aid of a Microsoft excel spread sheet where the values entered were; Range, Specific fuel Consumption of the Engine, Relative temperature, Mach Number, Aspect Ratio, Pressure at Cruise Altitude, Length of Fuselage and the number of Passenger including the Pilot. The take-off weight calculated according to the spread sheet was 44.03262kn. The spreadsheet also calculated the ratio of the operational empty weight to the takeoff weight, the ratio of payload weight to the takeoff weight and the fuel fraction. THRUST TO WEIGHT AND WING LOADING: Wing loading and thrust weight were calculated by individual group members and the best values which suited the type of aircraft in this project comparing to existing aircrafts were decided upon in group meetings. My individual calculations of TW & WS involved three different flight conditions (cruise – takeoff- landing) are shown below with graphs: Wing Loading and Thrust to Weight Determination for Jet Aircraft Cruise altitude/speed model Insert V and sigma(relative density)corresponding to cruise speed and altitude; use normal or fast cruise speed Cruise Conditions sigma = 0.32 CD0 = 0.02 AR = 6 k = 0.063654 V = (m/s) 231.3014 M = 0.7567 Ft = 0.7 rho SL = 1.225 Take-off Conditions mu = 0.035 Climax = 1.8 xg = (m) 500 rho = 1.225 CD0 = 0.1 Fig. 5: Wing Loading and Thrust to Weight Determination for Jet Aircraft Fig. 6: T/W vs W/S Fig 7: Landing Distance vs. W/S WING DESIGN: Starting this task of the project was not difficult as all the required calculations and formulas were provided in the hand out on wing design, I did start this task by calculating the wing gross area that affects the manoeuvrability of the aircraft. The initial factors to be decided upon were the wing loading (W/S) that in turn determined the surface area of the wing for a given MTOW. This was set by the graphs on landing distance vs. W/s and T/W vs. W/S, as shown in T/W vs. W/S spread sheet section which produced an initial W/S value of 3400 . This value estimated of W/S is small enough for project aircraft and suitable for cruise conditions at maximum take-off weights, thus avoiding more drag than is necessary during the cruise. From the wing loading of 3400, I calculated the wing area by dividing the take-off weight of 44032.62 KN by 3400 to give wing area of 12.9508 . And the aspect ratio was chosen by the author as 8. From the aspect ratio and wing area, I calculated wing span (b) using the formula AR = / S, in this case (b) worked out to be 8.815. With the wing span and area calculations, the tip and root chord lengths CT and CR respectively could be determined. The values of the tip and root chords are clearly calculated later on in this task. Using published data on critical Mach number analysis (cos = 0.7/M), given our preferred cruise speed of M = 0.7567, the sweep angle is given as cos = 0.7/0.7567 = = HIGH LIFT DEVICES: Aircrafts are fitted with high-lift device mechanisms that enable the aircraft to increase its flying capability during the flight operation. These devices are usually installed on the aircraft and are utilized when necessary especially during take-offs and landings. High-lift devices are categorized into powered and un-powered devices that commonly include flaps, leading edge extensions and slots (Aeromodelling). The flap is an adjustable portion of the aircraft’s rear wing that regulates the airflow hence increasing or reducing the aircraft’s lift. The slot on the other hand is located at the front or back of the wing and adjusts the airflow to the wing hence enabling the wing function more effectively. Powered high-lift devices influence the air flow on the wing’s exterior by using airflow emanating from the engine hence modifying the flap’s function. However, high-lift devices increase the aircraft’s drag during the flight by reducing its flight speed. Although these devices enable the aircraft increase or reduce its flying altitude, the resultant drag increases fuel consumption. In designing these devices, their overall impact on the aircraft are calculated since they generate a number of operational effects on the aircraft. Pressure distributions around the aircraft during flight operations determine greatly its manoeuvrability and fuel consumption. An aircraft’s runway requirements are highly dependent on the maximum lift and drag at high-lift conditions (Aeromodelling). The lift coefficient of an aerofoil with a trailing edge flap is given by: Where, Lift curve slope due to incidence Lift curve slope due to flap deflection Aerofoil incidence Zero lift incidences Flap deflection The value of may be found by using the table below, which gives the ratio of the flap chord to the aerofoil chord corresponding to the ratio of the lift curve slope due to incidence to lift curve slope due to deflection.  /  0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 0.282 0.396 0.481 0.55 0.609 0.661 0.707 0.748 0.785 0.818  /  = 0.25 (0.25 is quarter aerofoil section) So,  /  = 0.609 The aerofoil flap chord (c) = GMC = 1.4692 Flap chord (= (= 0.25 × 1.4692 (= 0.3673 m  =  × 0.609  = 0.1  = 0.1 × 0.609 = 0.0609 The increase in lift coefficient due to flaps, ΔCL, may be written: Where the co-efficient is the flaps effectiveness parameter, and it was found by using the table below: Flap Deflection 0 10 20 30 40 50 60 70 Effectiveness Parameter 0.7 0.69 0.65 0.58 0.49 0.41 0.32 - Using a flap deflection () of corresponding to an effectiveness parameter () of 0.69: = 0.69 × 0.1 × 0.609 × 10 = 0.42 The increase in lift coefficient due to flaps is not as large as the value suggests and is about of the value thus, = 0.42 × = 0.28 Form the equation We add the value as 0.28 and as 1.7 (from Theory of Wing section graph): = 1.7 + 0.28 = 1.98 Calculate the maximum lift coefficient of the whole wing () by using the equation below: () is the wing area ahead of the flaps but not including the flaps and it’s calculated as shown below:   So,  1.77 From Trust to Weight and Wing Loading, () was 1.8, and now it is 1.77. Therefore, the maximum lift coefficient of the aerofoil increases by 0.03. References Aeromodelling, High Lift Devices, Accessed on Dec. 09, 2009 from < http://www.myaeromodelling.com/wp/high-lift-devices >. Fielding, PJ 1999, Intro. to Aircraft Design, Cambridge: CUP, Retrieved from < http://assets.cambridge.org/052144/3199/sample/0521443199WSC00.pdf > on Dec. 11, 2009 Read More