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2 Defining the Design Domain The first estimation of aircraft weight used the iterative weight fraction method outlined in Roskam. This method calculates the weight fraction for each mission segment. The initial takeoff weight was guessed and the resulting empty weight was calculated. The resulting weight generally yields an unrealistic total weight fraction. In order to check the realism of the aircrafts weight fraction, a database of aircraft similar in mission was compiled. The aircraft gross takeoff weight was iterated until the total weight fraction landed on the historical trend. This trend is shown in Figure 2.1. The weights generated by this method include gross takeoff weight, fuel weight, and landing weight. The results of the weight fraction method are shown in Table 2.I. The assumptions made to generate the weight fractions are shown in Table 2.II. Vendetta Estimated Weight Fractions Cruise Back 16% Reserve 6% Misc. 5% Warm -up & Takeoff 6% Initial Climb 11% Dash Back 14% Dash Out 17% Cruise Out 25% Figure 2.1 - Historical Weight Fractions Table 2.I - Weight Fractions & Weights Mission Segment Weights Weights Start/Takeoff 6% Takeoff 108,400 lb (49,169 kg) Climb To Cruise 11% Empty 51,600 lb (23,405 kg) Cruise-Out 25% Fuel 47,600 lb (21,591 kg) Dash-out 17% Payload 9,054 lb (4,107 kg) Dash-Back 14% Fuel Weight Fraction 47.6% Cruise-Back 16% 45 Minute Reserve 6% Misc. 5% Total 100% 7
Many of the equations buried in the weight fraction method depend on assumed parameters. In many cases, values were assumed using figures and tables from Roskam, Nicolai, and Raymer. The assumptions used are listed in Table 2.2. The weight fraction method is by no means an accurate method for initial sizing. Inaccuracies of up to 10% are possible depending on the quality of the initial assumptions, and 20% is not uncommon for unusual missions such as the one outlined in the RFP. The weight fraction method may not be accurate for this type of aircraft due to the lack of similar aircraft in the database. There are really only three supercruise aircraft, the SR-71, YF-23, and the F-22. These aircraft all have a vastly different mission and may yield invalid sizes for the interdictor. Though the method may be flawed, it was used anyway due to the lack of a better method. Weight fractions provide a starting point for the weight of a proposed aircraft; however, the physical dimensions are not predicted. In order to determine the physical size, constraint plots were created. A constraint plot examines the relationship between two variables based on given requirements. Generally, the two variables used are wing loading and thrust to weight ratio. The RFP gives many constraints as shown earlier in Table 1.II. The majority of constraints can be written as functions of wing loading and thrust-to-weight ratio. This is the reason that it is a popular type of constraint plot. The equations for range, takeoff distance, and many others were found in Roskam, Nicolai, and Raymer. Many more assumptions were made to create the constraint plot; these are shown in Table 2.III. Table 2.II - Weight Fraction Assumptions SFC _Cruise 1.11 SFC _Dash 1.11 SFC _Turn 1.11 SFC _Loiter 0.8 L/D Cruise 10 L/D Dash 10 L/D Turn 10 L/D Loiter 12 Table 2.III - Constraint Assumptions C Lmax_TO 1.8 C Lmax_CR 1.2 C LCruise 0.2 AR 3 e 0.8 The constraint equations show how thrust to weight ratio and wing loading relate to a given performance constraint. This allows engineers to determine which combinations of thrust to weight ratio and wing loading are acceptable. The constraint plot for the RFP is shown in Figure 2.2. Note that any design point on the shaded side of a line would not meet the design requirements. This again depends on the accuracy of the aforementioned assumptions. The constraint plot clearly identifies a design domain, in which, any combination of thrust-to-weight ratio and wing loading will satisfy the design requirements. Combined with the weight fraction method, the constraint plot shows the physical size of the airplane. Because a preliminary weight was determined from the weight fraction method, the 8
Page 1 and 2: SAWE Paper No. 3232 Category No. 10
Page 3 and 4: Table of Contents Abstract_________
Page 5 and 6: List of Figures Figure 1.1 - Design
Page 7 and 8: Figure 13.4 - Rectilinear Vision Pl
Page 9 and 10: Nomenclature AIAA American Institut
Page 11 and 12: α Angle of Attack, degrees, rad.
Page 13 and 14: Table 1.II - Summary of Design Requ
Page 15 and 16: Figure 1.3 - F-15 Strike Eagle F-11
Page 17: Table 1.III - Comparison of the F-1
Page 21 and 22: 3 Configuration The current configu
Page 23 and 24: Another problem arises from the 20
Page 25 and 26: • Span = 54.7 ft • m.a.c. = 32
Page 27 and 28: 4 Stealth Considerations As mention
Page 29 and 30: Table 4.I - Common Ground Radars Ra
Page 31 and 32: 30 20 10 0 -10 -20 -30 -40 -50 1 GH
Page 33 and 34: 5 Aerodynamics The major aerodynami
Page 35 and 36: 1 GHz. 40º LE Sweep 10 GHz. 40º L
Page 37 and 38: 5.3 Wing Thickness The effect of wi
Page 39 and 40: 5.4 Airfoil The NACA 65A-003 airfoi
Page 41 and 42: 1 Section Lift Coefficient 0.9 0.8
Page 43 and 44: 90 80 Wing Cross Sectional Area (ft
Page 45 and 46: 6 Propulsion In developing the prop
Page 47 and 48: Table 6.II - RFP Dimensions Compare
Page 49 and 50: 38 Figure 6.1 - VAATE Goals
Page 51 and 52: TSFC 3.0 2.8 2.6 2.4 2.2 2.0 1.8 1.
Page 53 and 54: Deflection Angle Analysis for Mach
Page 55 and 56: Figure 6.8 - Cost Association with
Page 57 and 58: The inlet has a boundary layer dive
Page 59 and 60: 7 Materials and Structure The overa
Page 61 and 62: Vertical Attachment Points Horizont
Page 63 and 64: 8 Landing Gear Landing Gear design
Page 65 and 66: steel. This means that more braking
Page 67 and 68: Table 9.II - Final Component Weight
Page 69 and 70:
throughout the 5-hour mission. Usin
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10 Stability and Control To initial
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difference in neutral point and cen
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40 30 20 10 0 -10 -20 -30 -40 -50 2
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The Vendetta configuration utilizes
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-0.6 -0.4 UNSTABLE NEUTRAL -0.2 Lif
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Validation of an aircraft design su
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consumption. The main aspects of th
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In addition to the performance requ
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70,000 Altitude (ft) 60,000 50,000
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11.4 Mission Requirements By comple
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Table 11.VIII - Landing Flight Prof
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11.7 Alternate Missions In addition
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12 Payload Weapon internal layout d
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13 Cockpit Cockpit Design began wit
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that could be used in landing and t
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14 Systems The systems of the Vende
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14.3 Fuel System Initial fuel syste
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15 Manufacturing Several manufactur
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16 Cost Analysis The final and perh
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Appendix Threats Chart Common Surfa
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Wing Break Point Fuel Aft Fuselage
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The Vendetta Design Team Chris Atki
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References 1. “Ejection Systems,
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35. www.dfrc.nasa.gov/PAO/PAIS/HTML
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