AV I AV I S
SHORT RANGE,HIGH CAPACITY AIRCRAFT DESIGN UNIVERSITY OF CALIFORNIA,DAV I S
Design Report
June 10, 2020
Authors: Nahom Benyam Emre Mengi Christopher Pribilo Zachary Price Yashdeep Sidana Faculty Advisors: C.P. ”Case” van Dam, PhD Jared Sagaga Ryan Han 2020 AIAA Design Competition
Team Members AIAA Number Signature
Nahom Benyam 1109268
Emre Mengi 984297
Christopher Pribilo 1082933
Zachary Price 1109176 zachhim Yashdeep Sidana 1109292
1 Abstract
This report includes a detailed design of a short-range high capacity aircraft as well as a post-design analysis in accordance with the research for proposal given by the American Institute of Aeronautics and Astronautics. The report includes the work of the AVIAVIS student design team at the University of California, Davis who designed the
SRHC-530G aircraft. The designed aircraft aims to reduce the problem of overcrowded airports at major economic hubs without the size and cost that comes with long range capability. The SRHC-530G includes a twin-engine layout mounted on the back of the fuselage with a cruciform tail configuration, ensuring structural stability and clearance for the large GE9x engine nacelles. The aircraft is designed to carry 414 passengers in a two-class configuration with a twin-aisle setup and is optimized for short-haul routes of 700 nautical miles. The maximum range of the aircraft is
3700 nautical miles with maximum payload at 35,000 ft cruise altitude and cruising speed of Mach 0.8. The maximum take-off weight of the aircraft is 566,000 pounds with a take-off ground roll of 7500 feet to clear a 35 foot obstacle.
The SRHC-530G will also utilize hydrogen as a fuel source to reduce the carbon footprint, revolutionize air travel, and save money on operational costs. With its capabilities, the SRHC-530G is expected to lead the commercial aviation market in efficiency and reducing congestion at airports around the world.
3 Contents
1 Detailed Design - Hydrogen Storage 1
2 Introduction 12 2.1 Report Outline ...... 13
3 Concept Evolution & Final Configuration 14 3.1 Design Goals ...... 14
3.2 Market Study ...... 15
3.3 Baseline Aircraft Specifications ...... 17
3.4 Preliminary Sizing ...... 18
3.4.1 Weight Estimation ...... 18
3.4.2 Stall Speed, Takeoff, & Landing Sizing ...... 18
3.4.3 Preliminary Drag Polar ...... 19
3.4.4 Climb Sizing ...... 19
3.4.5 Sizing Diagram ...... 21
3.5 Carpet Plots ...... 23
3.6 Trade Study ...... 25
3.6.1 Fuselage Configuration ...... 25
3.6.2 Wing Location [33] ...... 26
3.6.3 Tail Configuration [30] ...... 27
3.7 Chosen Design Parameters ...... 30
3.7.1 Fuselage ...... 30
3.7.2 Engine Placement ...... 30
3.7.3 Wing Placement ...... 30
3.7.4 Tail Configuration ...... 30
3.8 Final Configuration ...... 31
3.8.1 3-D View of the Final Design ...... 31
3.8.2 Summary of the Aircraft Parameters ...... 34
4 Propulsion Systems 35 4.1 Propulsion Requirements & Engine Selection ...... 35
4.2 Alternative Fuel Source Candidate - Hydrogen ...... 36
4 5 Weight and Balance & Component Weights 37
6 Aerodynamics 40 6.1 Aerodynamics Overview ...... 40
6.2 Airfoil Selection and Characteristics ...... 43
6.3 High-Lift Systems ...... 45
6.4 Airplane Drag Breakdown ...... 49
6.5 CFD Analysis ...... 50
6.5.1 Wingtip Cant Angle Trade Studies ...... 51
6.5.2 Airplane (3D) Lift Curves (cruise, takeoff, landing) ...... 52
6.5.3 Airplane Drag Polars (Cruise, Takeoff, Landing) ...... 53
7 Stability and Control 54 7.1 Horizontal Tail Sizing ...... 54
7.2 Vertical Tail Sizing ...... 55
7.3 Aircraft Trim Diagram ...... 56
7.4 Control Surface Sizing ...... 58
7.5 Pitch, Roll, and Yaw Characteristics ...... 59
8 Performance 60 8.1 Takeoff and Landing ...... 60
8.1.1 Takeoff ...... 60
8.1.2 Landing ...... 60
8.2 Climb Requirements ...... 61
8.3 Cruise Performance & Payload-Range ...... 62
8.4 Comparison with Baseline/Competing Aircraft ...... 63
8.5 Comparison with Baseline/Competing Aircraft ...... 63
9 Structural Layout, Materials, Manufacturing 64 9.1 Aircraft Structure, Material, and Manufacturing Choices ...... 64
9.2 V-n Diagram ...... 66
10 Systems 67 10.1 Cabin Layout ...... 67
10.2 Cockpit ...... 70
5 10.3 Emergency Egress, Fire Protection Safety Systems ...... 71
10.4 Landing Gear ...... 73
11 Detailed Design and Analysis: Hydrogen Storage 75 11.1 Sizing and Structures ...... 76
11.1.1 Shell ...... 76
11.1.2 Frame ...... 76
11.2 Storage ...... 79
11.3 Materials and Insulation ...... 81
11.3.1 Materials ...... 81
11.3.2 Insulation ...... 81
11.4 Structural Analysis ...... 83
11.5 Conclusions and Future Work ...... 84
12 Cost and Utilization 85 12.1 Capital Costs ...... 85
12.2 Utility Cost ...... 87
12.3 Comparisons to Competing Aircraft ...... 90
12.4 Lifecycle CO2 Analysis ...... 90
13 Conclusions and Recommendations 91
14 References 92
15 Appendix 97 15.1 Fuselage Calculations ...... 97
15.2 Weight Excursion Diagram Calculations ...... 99
15.3 Payload-Range Diagram Calculations ...... 100
15.4 Drag Breakdown: Parasitic Drag ...... 102
15.5 Aircraft Part Sizing Based on Max Takeoff Weight ...... 103
15.6 Calculating Various Landing Gear Parameters ...... 104
List of Tables
1 AVIAVIS Design Specifications ...... 12
6 2 Comparable Aircraft Specifications [3][6][8][9][19] ...... 17
3 Preliminary Sizing Results ...... 22
4 Summary of Aircraft Parameters: Aerodynamics ...... 34
5 Summary of Aircraft Parameters: Velocities ...... 34
6 Summary of Aircraft Parameters: Fuselage Specs and Weights ...... 34
7 Summary of Aircraft Parameters: Engine Specifications ...... 34
8 Engine Selection Trade Study [55] ...... 35
9 Initial Weight Breakdown of SRHC-530G ...... 37
10 Aerodynamic Requirements ...... 40
11 Wing Geometrical Parameters ...... 40
12 Airfoil Trade Study (Re 40million) ...... 43 ⇡ 13 Drag Breakdown for Cruise (Open VSP) ...... 49
14 Wingtip Cant Angle Trade Study for Cruise Condition ...... 51
15 Parameters for Aircraft Trim Calculations ...... 56
16 Maximum Control Surface Deflection Angles ...... 59
17 Takeoff Performance Analysis Results ...... 60
18 Landing Performance Analysis Results ...... 60
19 Comparison of SRHC-530G with Competing Aircraft [6] ...... 63
20 Comparison of SRHC-530G with Mission Specs ...... 63
21 Composite Diagram Data ...... 66
22 Cabin Layout Specifications ...... 67
23 Galley, Lavatory, and Emergency Exit Specifications ...... 67
24 Fuselage Dimensions ...... 69
25 ULD Specifications ...... 69
26 Tank and Fuselage Dimensions ...... 76
27 Hydrogen Tank Frame Dimensions ...... 77
28 Properties of Select Materials[64] ...... 81
29 Hydrogen Tank FEA Results ...... 84
30 Labor Costs ...... 85
31 Annual Acquisition and Development Price in millions ($) ...... 85
32 Average Direct Operating Cost per Block Hour for Widebody Aircraft Adjusted for 2020 US Dollars
[52] ...... 87
33 Direct Operating Cost analysis for 700 nm and 2500 nm reference missions ...... 89
7 34 Annual Direct Operating Cost Analysis ...... 90
35 Unit Cost of comparable Aircraft [55] ...... 90
List of Figures
1 Mission Profile for SRHC-530G ...... 14
2 Aircraft Operating Costs per Seat-Mile for Different Stage Lengths [14] ...... 15
3 737-800NG [15] ...... 16
4 Boeing 777-300ER [16] ...... 16
5 Boeing 777-200 [17] ...... 17
6 Preliminary Drag Polars ...... 20
7 Preliminary Sizing Diagram ...... 21
8 Design space for the Short Range High Capacity Aircraft at 566,000 MTOW at a cruise altitude of
35,000ft (SA) ...... 23
9 Design space for the Short Range High Capacity Aircraft at 566,000 MTOW at a cruise altitude of
35,000ft (SA) and Ma = 0.80 ...... 24
10 Fuselage Configuration [23] ...... 25
11 Economy Class Fuselage Cross-Section ...... 26
12 Wing Location Configuration [33] ...... 27
13 TopView ...... 31
14 Side View ...... 32
15 Front View ...... 32
16 Isometric View ...... 33
17 GE9x ...... 36
18 Required Static Margins for Different Aircraft Models [11] ...... 38
19 Weight Excursion Diagram for SRHC-530G ...... 38
20 Wing Planform ...... 41
21 Profile Drag Difference During Climb and Cruise from Traditional Winglets ...... 42
22 Tentative Flight Plan for Different Wingtip Angles ...... 42
23 Boeing Airfoil J Geometry ...... 43
24 Lift Curve at Cruise for Boeing Airfoil J (Re 40million) ...... 44 ⇡ 25 Drag Polar at Cruise for Boeing Airfoil J (Re 40million) ...... 44 ⇡ 26 Definition of Flapped Wing Area [12] ...... 46
8 27 Sectional effectiveness for single-slotted flaps and Fowler flaps [40] ...... 46
28 Airplane Lift Curve At Landing Conditions with different Flap Deflections ...... 47
29 Diagram of Krueger Flaps and Slats [28] ...... 48
30 Flow Separation during Cruise ...... 50
31 3D Lift Curve for Take-off and Landing, Wingtip Angle = 45 ...... 52
32 3D Lift Curve for Cruise, Wingtip Angle = 15 ...... 52
33 3D Drag Polar for Take-off and Landing, Wingtip Angle = 45 ...... 53
34 3D Drag Polar for Cruise, Wingtip Angle = 15 ...... 53 35 Horizontal Tail Sizing and Static Margin for a Takeoff Weight of 566,000 lbs ...... 55
36 Vertical Tail Weight Sizing ...... 55
37 Aircraft Trim Diagram for Cruise Conditions at W = 566,000 lbs and a CG of 0.857 of MAC . . . . . 57
38 Thrust Required Vs Thrust Available ...... 61
39 Payload-Range Diagram for SRHC-530G ...... 62
40 Aircraft Material Selection ...... 65
41 Composite V-n & Gust Diagram ...... 66
42 Economy Class Fuselage Cross-Section ...... 68
43 Business Class Fuselage Cross-Section ...... 68
44 Cabin Layout ...... 69
45 Honeywell AIMS [47] ...... 70
46 Landing Gear Configuration ...... 74
47 Hydrogen Tank Shell - Iso View ...... 77
48 Hydrogen Tank Frame - Iso View ...... 77
49 Hydrogen Tank Frame - Side View ...... 78
50 Hydrogen Tank Frame - Top View ...... 78
51 Hydrogen Density at Various Pressures and Temperatures [58] ...... 79
52 Cryogenic Temperature, Pressure, and Density Values [59] ...... 80
53 Fuselage and Tank Layers ...... 82
54 Placement of Insulation and Liners ...... 82
55 Tank Frame FEA Stress Heat Map and Deformation Plot ...... 83
56 Cost Per Unit Adjusted for Inflation with 15% Profit ...... 86
57 Variable Direct Operating Cost per Block Hour ...... 88
58 Fixed Direct Operating Cost per Block Hour ...... 88
59 Total Direct Operating Cost per Block Hour ...... 89
9 60 Payload Range Diagram Example ...... 101
61 OpenVSP Parasite Drag Calculation ...... 102
62 Aircraft Sizing as Percentage of MTOW [21] ...... 103
63 Lateral Tip-Over Criterion (Top View) ...... 104
64 Lateral Tip-Over Criterion (with Inclined Angle) ...... 105
65 Ground Clearance Angle ...... 106
10 Nomenclature
Abbreviations Symbols AC Aerodynamic Center m˙ Mass Flow Rate CG Center of Gravity b Wingspan CFD Computational Fluid Dynamics CD Drag Coefficient FAR Federal Aviation Regulation Cl Lift Coefficient MAC Mean Aerodynamic Chord CM Pitching Moment Coefficient RFP Request for Proposal D Diameter ULD Unit Loading Device df Fuselage Diameter l location lf Fuselage Length P Power, Loading T/WTO Thrust-to-Weight Ratio S Wing Area v Velocity W Weight WTO/S Takeoff Wing Loading Subscripts TO Take Off max Maximum n Nose Gear m Main Gear
11 2 Introduction
As international commerce matures, the need for efficient short range travel between various economic hubs is on the rise. The absence of a high capacity – short range transport jet is causing congestion problems at major airports around the world. According to the Department of Transportation, delays are the most common passenger complaint, accounting for about 40% of all complaints [1] . Furthermore, the growth of upcoming economic hubs like China,
India, and Nigeria will contribute more to this already increasing congestion problem. This, in conjunction with de- pleting fuel reserves and unstable oil prices creates an unsteady environment for both airliners and travelers alike making the efficient allocation of airspace of paramount importance.
The problem associated with modern airliners lies in the absence of enough seating to meet both consumer and environ- mental demands. Most of the aircraft used in medium - short haul flights were designed at a time where the demands of air travel were not as persistent as those of today. Data from the world bank shows that in 1970, international demand for air travel was 310 million users per year, today this number is a massive 4.25 billion [2]. Furthermore this number is predicted to grow by 5.2% per annum, for the next 20 years [2]. The number of active fleet will rise as a result, exacerbating the congestion issue. The proposed aircraft is designed to mitigate this problem by doubling the passenger capacity and revitalizing the propulsion system, targeting both environmental and consumer concerns.
The proposed propulsion system will include a high bypass ratio (10:1) engine with the capability to run hydrogen fuel to reduce operational costs as well as reduce greenhouse gas emissions. The fuel capacity for current short range aircraft’s like the Boeing 737 carry a maximum of 230 passengers with a 5,970 gallon fuel tank. With the assumption of $3 a gallon for fuel and $3 per gallon for carbon tax, the cost of operation is $34570 per full tank. The current unit cost of hydrogen is 14 per kg which is equivalent to $5.60 per gallon of fuel — furthermore it is predicted that the pr kg cost of hydrogen will fall to $8 on the lower end. [3] This translates to roughly $3.2 per gallon of equivalent energy, assuming the same volume for the fuel tank, the cost of operation per fuel tank per flight is roughly $18,528, with no carbon tax due to the ‘clean’ properties of hydrogen propulsion. This 46% savings in operative costs translate to a higher profit margin for airlines while also decreasing contribution to greenhouse gasses.
Table 1: AVIAVIS Design Specifications
Aviavis High Capacity Transport Jet Specifications Min No. Passengers 400 Cruise Speed Mach 0.80 Cruise Altitude 35,000 ft Range 3500 (n.mi) Maximum Take-off length 9000 ft
12 The high capacity short range transport jet outlined in this report will comply with CFR Title 14- Part 25 regulations
[4] with a service entry date of 2029. The main requirements for the aircraft are outlined in Table 1.
Key technologies would need to be developed to ensure that hydrogen usage is feasible. Hydrogen must be stored in highly pressurized tanks to preserve its liquid state, thus fuel tank technology needs to be developed to handle these conditions. Furthermore as hydrogen is highly combustible, fire prevention systems as well as sophisticated fuel dump mechanisms need to be developed. Lastly, improvements on existing heat exchangers need to be made to ensure that they can accommodate the safe and efficient heating of hydrogen fuel.
2.1 Report Outline
The introductory section of the report details the the issues modern airliners face, the RFP for this issue, as well as the design choices made to address those issues. The following section includes a detailed trade off analysis for major components of the aircraft and the systems level design choices made by the team. Current competing aircraft are also analyzed to ensure the proposed design further improves on what is currently available. Following this an engine matching study is conducted with the rendered design to meet and exceed performance under provided conditions. A weight analysis is then conducted to determine the placement of the wings and landing system. The aerodynamics of the finalized configuration is then analyzed and iterated to minimize the contribution of drag during cruise flight.
Lastly, the performance of the aircraft at take off, cruise, and landing is analyzed to ensure the RFP demands are met.
Conclusions and future recommendations follow.
13 3 Concept Evolution & Final Configuration
3.1 Design Goals
The RFP provided by the AIAA entails that the aircraft should be designed for an entry to service of 2029, with a seating capacity of 400 and max range of 3500 nautical miles. However, the aircraft is to be optimized for a reference mission of 700 nautical miles, to mimic the common distance between large economic hubs. The primary design goals of this aircraft is to produce a cost efficient aircraft while maintaining or improving reliability and repairability. The mission specifications for the requested aircraft design by AIAA are briefly tabulated in Table 1. The expected mission profile for the aircraft is shown below.
Figure 1: Mission Profile for SRHC-530G
14 3.2 Market Study
Before designing the aircraft with regard to the mission specifications provided, it is important to review the existing commercial airplane market segments relevant to mission profile of the aircraft. It is seen that the two main markets segments for commercial aviation are short-haul airplanes with low capacity and long-range airplanes with high capac- ity. Airlines’ demand from the aircraft manufacturers is to reduce the cost of operation of the airplane for the selected route. A study from Wenbin and Hansen [14] shows that optimal number of seats for cost per seat-mile decreases as the route length decreases, as seen in Figure 2. Therefore, it is not surprising to not see many aircraft in mid-segment.
Figure 2: Aircraft Operating Costs per Seat-Mile for Different Stage Lengths [14]
However, it can be hypothesized that the consideration of the direct operating cost of the aircraft during the aircraft design process does not address the problems that exist beyond the operation of a sole aircraft. As more airlines opt in for short range low capacity aircraft to have frequent flights to attract more customers, the existing airports simply cannot keep up with the fleet expansion of the airlines, causing overcrowding of the airports. To solve such problems, the two main market segments mentioned previously should be discussed.
An example of an aircraft from the short-haul low capacity segment is the Boeing 737-800, which satisfies the needs for frequent flights offered by the airlines. With the increase in use of air travel, the corresponding short range low
15 capacity aircraft fleet expansion of the airlines lead to overcrowded airports problem mentioned previously.
Figure 3: 737-800NG [15]
The other market segment is the long range, high capacity aircraft that operate in less frequent routes while carrying a large amount of passengers to balance the operating cost of the aircraft. An example of an aircraft for this market segment is the Boeing 777 series. The required fuel to fly long distances along with a heavier payload leads to more complex lift systems and more powerful engines equipped on the aircraft. Therefore, it is important to consider the complexity and the weight penalty of the systems and engines to be installed on the proposed aircraft design, in order to reduce the operating costs of the aircraft and make it a feasible solution to the problems listed.
Figure 4: Boeing 777-300ER [16]
16 3.3 Baseline Aircraft Specifications
The initial sizing of the aircraft in design requires a baseline aircraft to base the initial aircraft parameter assumptions, such as take-off weight, empty operating weight, etc. The possible candidates for the baseline aircraft were McDonnell
Douglas DC-10, due to its comparable range to the RFP specifications and 270 passenger capacity, Boeing 747-400ER, due to its high capacity of 568 passengers, Boeing 777-300ER, due to its 451 passenger capacity, Boeing 777-200, due to its shorter range compared to the other aircraft with 400 passenger capacity. The passenger capacity given are for 2-class configurations. The specifications for these aircraft are given in Table 2 below.
Table 2: Comparable Aircraft Specifications [3][6][8][9][19]
Aircraft Cruise Speed Range Cruise Altitude Passengers Length MTOW DC-10 490 kt 4000 nmi 42,000 ft 270 182.1 ft 580,000 lb 747-400ER 507 kt 7670 nmi 45,000 ft 568 231.1 ft 910,000 lb 777-300ER 490 kt 7930 nmi 35,000 ft 451 242.4 ft 775,000 lb 777-200 490 kt 5240 nmi 35,000 ft 400 209.1 ft 545,000 lb
The table shows that the best option for the baseline aircraft among the selected set is Boeing 777-200. This aircraft
fits the specified passenger count of 400 in 2-class configuration while maintaining a relatively short range compared to the other high capacity aircraft. Therefore, the following sections use the specifications of this aircraft for initial guesses for preliminary sizing that was used to configure the designed aircraft.
Figure 5: Boeing 777-200 [17]
17 3.4 Preliminary Sizing
3.4.1 Weight Estimation
The take-off weight of the aircraft is the combined weight of payload, fuel, and aircraft operating empty weight. The payload weight is passenger and baggage weight, calculated by assuming 200lbs and 30lbs for 400 passengers respec- tively as required by mission specifications. The fuel weight is calculated by estimating fuel percentages leftover after each flight phase to judge how much fuel is consumed as a fraction of take-off weight for the duration of the entire
flight, including 1 hour loiter. Fuel reserves for 30 minutes of extra flight at cruise speed is also added to the fuel weight. The percentages are obtained from Roskam Part I [10].
Lastly, the operating empty weight is a combination of crew weight, trapped fuel and oil weight and empty structural weight. The crew is assumed to include 10 people with the same weight assumptions as passengers. The number of crew members are specified in the AIAA RFP as a tradable requirement: 2 pilots and 8 crew members. The trapped fuel weight is a small fraction of the empty operating weight and empty structural weight is derived from the compar- ative Boeing 777-200 aircraft.
This method of calculating take-off weight makes the expression implicit so an initial take-off weight is estimated and calculations are iterated till convergence. The obtained value after convergence analysis yielded a take-off weight of
WTO = 566, 000lbs.
3.4.2 Stall Speed, Takeoff, & Landing Sizing
Stall speed is an important parameter in aircraft performance and design process. Given that the aircraft has a spe- cific maximum take-off weight, a wing reference area comparable to similar aircraft, given air density at a reference
altitude, and CLmax obtained from Roskam Part I, the stall speed can be chosen to obtain a wing loading limit for the aircraft. Although FAR 25 requirements do not enforce a stall speed range, a stall speed of Vstall = 109knots was chosen. This value is comparable to the aircraft in the same segment of wide-body aircraft with high passenger capacity.
Next, an empirical expression from Roskam is used to determine the size of the airplane wing and powerplant require- ments. FAR Part 25 requires the distance to the point the airplane is 35 ft above takeoff surface, this point is calculated using reference Roskam Part I as well. Using the predetermined maximum lift coefficient, density ratio, and wing loading, the corresponding thrust to weight ratio was found.
18 Similar to takeoff requirements, FAR 25 requires the landing distance to be measured from the point where the aircraft is 50 feet above the landing surface. Roskam Part I provides empirical expressions for the aircraft landing distance requirements, which include the aircraft surface area, acceleration, and landing velocity. Typically, the weight per aircraft wing area at landing is less than or equal to the weight per aircraft wing area at take off.
3.4.3 Preliminary Drag Polar
A preliminary drag polar is obtained for different flight phases by estimating the wetted surface area and flat plate area of the aircraft, using Preliminary Sizing Airplanes by Jan Roskam. For a standard transport jet and previously calculated take-off weight, the wetted surface area is found to be 22,508 ft2. Furthermore, a jet transport average of
2 equivalent Cf of 0.003 is assumed and a flat plate area of 67.52 ft is calculated. Profile drag coefficient CD0 is found to be 0.0146 and induced drag for different flight phases is then calculated using an aspect ratio of 11.53, and
Oswald span efficiency of 0.8 for cruise. The aspect ratio of the aircraft was determined from the initial configuration choices made on Open Vehicle Sketch Pad (OpenVSP) software by NASA [18], while Oswald span efficiency factor was chosen in accordance with the examples given in Roskam Part I.
The different phases portrayed on the drag polar include cruise, take-off with and without landing gear exposed, and landing with and without landing gear exposed. The drag coefficients and changes in span efficiency is obtained from
Roskam Part 1. The following drag polar curves were obtained.
3.4.4 Climb Sizing
As part of the Federal Aviation Regulations, aircraft need to fulfill certain climb requirements for the different stages of their flight. As a jet transport aircraft with a weight higher than 12,500 lbs, the aircraft in design was analyzed under
FAR 25 requirements. To fulfill the requirements, the needed thrust-to-weight ratio was calculated for the initial climb segment (FAR 25.111 - OEI), transition climb segment (FAR 25.121 - OEI), second climb segment (FAR 25.121 -
OEI), en-route climb (FAR 25.121 - OEI) requirements. These were calculated in accordance with the take-off weight and one engine inoperative condition. For the landing requirements, FAR 25.119 - AEO and FAR 25.121 - OEI were considered, which are known as go-around or balked landing requirements. The most strict thrust-to-weight ratio requirement was plotted on the sizing diagram.
19 Figure 6: Preliminary Drag Polars
20 3.4.5 Sizing Diagram
The FAR 25 requirements were used to size the aircraft for cruise, climbing, takeoff and landing under 9000ft and cor- responding stall and approach speeds. The thrust(in terms of weight) required for cruise against cruise drag, and the thrust required to climb with only one engine operative(for a second climb FAR 25.121) was calculated and plotted.
In addition, the wing-loading caused in taking-off within 9000ft, and wing loading caused by landing within 9000ft and corresponding approach speeds was calculated as well, and graphed with different coefficients of lift. These coef-
ficients are required to be achieved using high lift devices used in taking-off and landing.
Figure 7: Preliminary Sizing Diagram
21 On the sizing diagram, the point chosen for the designed aircraft, SRHC-530G, is marked by the blue marker. The location on the y-axis, thrust-to-weight ratio, was chosen accordingly to the proposed propulsion system of a twin- engine GE9X set-up. The thrust provided by the GE9X engines was found to be a combined 210,000 lbf while the take-off weight of the aircraft was used to get the maximum ratio of T/W [20]. The x location was chosen to obtain the maximum wing loading while maintaining operation away from the stall speed sizing requirement for W/S.
The resulting values were T/WTO =0.35 and WTO/S = 120psf, which yielded a maximum required thrust of T = 198, 100lbs, which is below the maximum thrust that can be supplied by two GE9X engines, while the wing area
2 was found to be Swing = 4717ft . Other parameters obtained through the preliminary sizing process is tabulated below:
Table 3: Preliminary Sizing Results
Parameter Value Take-off Weight 566,000 lbs Operating Empty Weight 296,000 lbs Empty Weight 294,170 lbs Fuel Weight 178,179 lbs Maximum Lift Coefficients Clean 0.51 Take-off 2.14 Landing 2.69 Profile Drag Coefficient 0.0146 Aspect Ratio 11.53 Take-off Wing Loading (WTO/S) 120 psf Wing Area 4717 ft2 Mean Aerodynamic Chord 20.23 ft Take-off Thrust-to-Weight Ratio (T/WTO) 0.35 Take-off Thrust 198,100 lbs
22 3.5 Carpet Plots
To determine the final design of the aircraft, several carpet plots were created [21]. These plots help choose certain design parameters for the designed aircraft, SRHC-530G.
Figure 8: Design space for the Short Range High Capacity Aircraft at 566,000 MTOW at a cruise altitude of 35,000ft (SA)
The first carpet plots shows the variation in wing loading and maximum lift-to-drag ratio of the aircraft for various aspect ratio and design Mach number values.The design point is determined to be an aspect ratio of 11.5 and cruise speed of Mach 0.8, which yields a maximum lift-to-drag ratio of 22. This increases the required wing loading to around 150 lb/ft2, which is over the stall speed sizing limit of 120 lb/ft2. Thus, CFD analysis will determine the
final flight parameters to increase the fidelity of results and ensure that the wing loading limit is not exceeded.
23 Figure 9: Design space for the Short Range High Capacity Aircraft at 566,000 MTOW at a cruise altitude of 35,000ft (SA) and Ma = 0.80
The next carpet plots shows the design space for the designed aircraft at 566,000 MTOW at a cruise altitude of 35,000ft and cruise speed of Mach 0.80. The variation in maximum lift-to-drag ratio and wing loading with varying effective aspect ratio (AR e) and zero-lift drag coefficient can be seen in the graph. For an effective aspect ratio of 9.2, ⇤ maximum lift-to-drag ratio of 26.5 can be achieved if profile drag coefficient can be reduced to 0.0100. Therefore, further aerodynamic analysis will help refine the aircraft design to minimize drag and achieve the theoretical maximum lift-to-drag ratio values obtained through the carpet plots.
24 3.6 Trade Study
3.6.1 Fuselage Configuration
From the Aircraft Performance Design lecture notes, Professor van Dam noted that the fuselage is becoming more ovular to reduce wetted area. Figure 10 was shown in lecture with different fuselage configurations.
Figure 10: Fuselage Configuration [23]
To determine the best fuselage shape for the designed aircraft, an initial seating layout is selected. The selected configuration is 3-3-3 for economy and 2-2-2 for business class. This layout is selected because of its popularity among airlines and the adequate space it provides for each passenger for both classes. In addition, the selected configuration is ideal for boarding time. Next, for the selected layout, a variety of cross-sectional shapes are considered in terms of aerodynamic performance and ease of operation for airliner procedures.
Cross-Section Shape Circular Oval Custom Cross-Section CD 0.21937 0.20262 0.20976 Wetted Area 12218 ft2 11333 ft2 11726 ft2 Cargo Capacity Largest Smallest Medium
25 When the drag count and wet area of the fuselage is considered, the oval shape performs the best because of its low drag characteristics. However, the oval shape has a smaller cargo capacity compared to the circular cross-section.
So, there should be compromise between the cargo space and low drag performance for the fuselage body. Because the aircraft is expected to be fitted with engines that uses hydrogen as fuel, the fuselage needs to accommodate heat exchangers to heat the hydrogen fuel. So, a custom fuselage cross-section is created, which is shown below:
Figure 11: Economy Class Fuselage Cross-Section
This cross-section is to be used for the aircraft to accommodate the extra cargo space required by the hydrogen fuel set-up while maintaining a lower drag count compared to a simple circular shape. The lower deck cross-sectional shape also allows for two LD3-45 unit loading devices to be fitted next to each other. As a result, it is expected that there will be more space available in the lower deck for the fuel storage and heat exchanger devices. In addition, compared to an oval shape, this custom cross-sectional shape will increase the passenger comfort due to the extra overhead space available.
3.6.2 Wing Location [33]
Mid-wing and Low-wing configurations were considered in the design of the SRHC. Below is a short description of each setup:
Mid-Wing Configuration:
The wings are exactly at the midline of the airplane. Mid-Wing airplanes are very well balanced and their design allows for a large control surface area (portions of the plane involved in steering). This configuration is very maneu-
26 verable but not is not as stable as high wing airplanes.
Low Wing Configuration:
The wings are below the midline of the airplane. This configuration is more stable than mid-wing airplanes but not as much as high wing airplanes. Although this is not a desired characteristic, low-wing airplanes are more maneuverable than high-wing airplanes. In addition, the lower height of the wings allows for easy engine maintenance if the engines are mounted under the wing.
Figure 12: Wing Location Configuration [33]
Furthermore, the low-wing configuration was chosen for the reasons listed above as well as the chosen fuselage design. With cargo in the lower deck, wings are best supported when attached to the lower deck, also allowing for the passengers to have the ability to look over the wing with an unobstructed view.
3.6.3 Tail Configuration [30]
The airplane’s tail design is crucial since it controls and stabilizes the airplane in both up-and-down movements of pitch and side-to-side movements of yaw.
Conventional Tail:
27 The Conventional configuration is used in the A300, B777, and B747. This design provides adequate stability and control with the lowest structural weight.
T-Tail:
The horizontal stabilizer is positioned at the top of the vertical stabilizer. This allows for the horizontal stabilizer to be above the propeller flow and wing wake. Since the horizontal stabilizer is more efficient, it can be made smaller and lighter. The placement of the horizontal stabilizer on top of the vertical stabilizer also makes the vertical stabilizer more aerodynamically efficient, which allows the size to be reduced.
Nevertheless, the T-tail layout imposes a bending and twisting load on the vertical stabilizer, which requires a stronger and heavier structure. This kind of load is avoided in the conventional configuration. At a high pitch angle associated with landing, the horizontal stabilizer of the T-tail may be immersed in the slower and more turbulent flow of the wing wake. The T-tail configuration is used in the B727, MD-90, and DC-9.
Cruciform-Tail
The Cruciform-Tail configuration is a mix between the conventional and T-tail designs. The horizontal stabilizer is up and away from the jet exhaust and wing wake. The lifting of the horizontal stabilizer exposes the lower part of the vertical stabilizer, as well as the rudder, to undisturbed airflow. Undisturbed airflow on the rudder is important for recovery from spins. The Cruciform-Tail is used in the North American Rockwell B-1B supersonic bomber and the
Dassault Falcon 100.
Dual-Tail:
The Dual-Tail design places two vertical stabilizers at the ends of the horizontal stabilizers. This design places the ver- tical stabilizers in the prop wash of wing-mounted propellers, which gives good directional control during low-speed operations. This also allows for a smaller, lighter, and more aerodynamically efficient horizontal stabilizer. However, the overall weight of the plane with a dual-tail design is greater than that of a plane with a conventional-tail design.
This configuration is used in the Republic Fairchild A-10, the Ercoupe, and the Consolidated B-24.
Triple-Tail:
This design has two vertical stabilizers placed at the ends of the horizontal stabilizers and one mounted on the fuselage.
28 This configuration is ideal for when the height of the vertical stabilizer must meet certain restrictions. The Lockheed
Constellation and the Grumman E-2 Hawkeye are examples of this design.
V-Tail:
The ideal advantage of the V-tail design is that the two surfaces may serve the same function as the three required in the conventional tail. This configuration would help reduce the drag of the tail surfaces and reduce the weight in the empennage. However, wind tunnel studies show that the V tail achieves the same degree of stability as a conventional tail design. The area of the V tail would have to be about the same size as the conventional tail. The V-tail design has maneuverability challenges due to the presence of adverse coupling. An example of this configuration includes the
Beech-craft Bonanza V-35.
Inverted Y-Tail:
The inverted Y-Tail design is a conventional tail with the outer ends of the horizontal stabilizers lower than the ends attached to the fuselage. This was used in the F-4 Phantom to keep the horizontal surfaces out of the wing wake at high angles of attack.
Twin-Tail:
The twin tail is a feature in various air superiority fighters used by the military. With two vertical stabilizers, the twin tail is more effective than the conventional single tail of the same height. This design is used in the F-14 Tomcat and the F/A-18 Hornet.
Engine Location:
The engines were placed at the rear of the fuselage, due to large diameter turbofans, as well as for the placement of future larger bypass ratio engines. Since the engines will not be placed on the wing of the plane, the lateral ground clearance criterion is avoided. This is key as the aircraft has a low-wing configuration. Rear mounted engines will be less accessible for maintenance and require longer fuel lines. However, this configuration allows for a smaller rudder, avoids a likely encounter with debris, and allows for a more aerodynamically efficient wing without interference from the engine. Further details can be found in Chapter 3.
29 3.7 Chosen Design Parameters
3.7.1 Fuselage
A twin aisle cabin layout with a non-circular fuselage was chosen. A twin aisle configuration allows for a smaller cross sectional diameter, which in turn reduces the overall drag experienced by the aircraft. Furthermore a non- circular configuration was chosen to remove unnecessary space underneath the fuselage as the mission specification doesn’t require the carrying of cargo along with passengers and baggage. This in turn reduces overall structural weight while also minimizing drag.
3.7.2 Engine Placement
The engines were placed at the rear of the fuselage, due to large diameter turbofans, as well as for the placement of future larger bypass ratio engines. Further details can be found in Chapter 3.
3.7.3 Wing Placement
A low-wing configuration was chosen as it is the industry standard with most transport aircraft having this configu- ration. Low-wing provides the structural components like the spars of the wing to go through the lower part of the fuselage. This makes integration with cargo and passengers optimal as the top half of the fuselage is preserved for passengers only. In addition, the engines can be serviced easily if they are mounted under the wing. Currently, the engines are placed near the empennage, but low-wing configuration would be beneficial if the engines are chosen to be moved under the wings. In that case, to account for clearance, more wing dihedral may be added.
3.7.4 Tail Configuration
A cruciform tail was chosen for the aircraft configuration due to its stability properties. The placement of the engines prevent a conventional horizontal tail, thus either a T-Tail or cruciform tail were the available options. However, during stall, the disturbed flow from the wake of the wing will flow over the T-tail and render the elevators ineffective.
This causes the airplane to enter super-stall, a condition that is to be avoided. One way to mitigate this would be to implement a stick pusher system into the aircraft, which rattles the stick/yoke when the system senses super-stall conditions. However, the implementation of a system like this causes opportunities for failure thus a cruciform tail configuration was chosen to negate all super stall effects.
30 3.8 Final Configuration
3.8.1 3-D View of the Final Design
The following three-view diagram shows the take-off and landing wingtip configuration.
Figure 13: Top View
31 Figure 14: Side View
Figure 15: Front View
32 Figure 16: Isometric View
33 3.8.2 Summary of the Aircraft Parameters
Table 4: Summary of Aircraft Parameters: Aerodynamics
Aerodynamic Parameters 2 Swing 4717 ft AR 11.5 MAC 20.23ft Taper Ratio 0.176 CLcruise 0.51 CLMax,TO 2.14 CLMax,L 2.69 CD0,cruise 0.0131 Flat Plate Area 67.98 ft2 Airfoil Beoing Airfoil J
Table 5: Summary of Aircraft Parameters: Velocities
Velocities Vcruise Mach 0.8 VTO 202.2 ft/s VLD 211.4 ft/s
Table 6: Summary of Aircraft Parameters: Fuselage Specs and Weights
Fuselage and Weight Parameters Capacity 414 (2-class) Length 239.8 ft Cabin Width 18.2 ft Fuselage Width 19.3 ft Empty Operating Weight 296,000 lbs Maximum Take-off Weight 566,000 lbs Maximum Payload Weight 92,000 lbs
Table 7: Summary of Aircraft Parameters: Engine Specifications
Engine Specifications GE9x Max Thrust [lbf] 100-105,000 Weight [lbf] 40,000 Length [in] 290 Diameter [in] 134 Pressure Ratio 60:1 Bypass Ratio 10:1
34 4 Propulsion Systems
4.1 Propulsion Requirements & Engine Selection
The thrust to weight parameter was computed using the sizing diagram populated in Figure 7 and was found to be
0.35. Using Roskam part one, the takeoff weight was determined to be 566,000 lbs, resulting in a thrust requirement of 198,100 lbf. A twin engine arrangement was desired for the configurations of this aircraft, requiring each engine to produce roughly 100,000 lbf pounds of thrust. A trade study of a range of high performance engines currently used in the market were analyzed, the table below summarizes findings:
Table 8: Engine Selection Trade Study [55]
Max Thrust Weight Diameter Engine Pressure Ratio Bypass Ratio [lbf] [lbf] [in] GE90 82 - 97,000 19,000 123 40:1 8-9:1 PW400 90-99,000 16,260 112 34-42:1 5.8-6.4:1 RR Trent 800 77- 93,00 14,000 110 34-40:1 6.4:1 GE9x 100-105,000 40,000 134 60:1 10:1
Per the requirements of the RFP, one of the primary considerations is to reduce overall operational costs for a reference distance of 700 nm. Furthermore as this aircraft has an entry into service year of 2029, stricter requirements for fuel efficiency and economic friendliness is a paramount parameter as the Advisory Council for Aviation Research and
Innovation in Europe (ACARE) has set ambitious goals for 2050. With an aim is to reduce fuel emissions by 75%,
NOx by 90% and noise by 60% relative to aircraft’s in 2000 . Alternative fuels are considered in a later section to further address this issue. From the analyzed engines it can be seen that the GE9X proves to be the best option.
The high pressure ratios raise the inlet temperature of the fuel, which consequently increases net thrust and decreases specific fuel consumption. The GE9x also has the highest bypass ratio on the table shown above, high bypass ratios ensure that large amounts of air are passed through narrow nozzles which increase thrust whilst using the same amount of fuel. According to GE, the combination of these two parameters provide a 10% improvement in fuel burn when compared to the GE90-115B, a significant decrease when considering roughly 19% of an airlines operational cost is fuel [20].
The downside of using the GE9x comes from the large fan diameters as well as the considerably heavy weight. Thus, to mitigate ground clearance issues and reduce the amount of debris entering the engines, it was decided that mounting the engines on the rear of the fuselage would be the best route. Another option would be to mount the engines atop the wings, structural implications of these two configurations are currently being explored. The weight of these engines were considered when computing the initial sizing and are included in the takeoff weight.
35 Figure 17: GE9x
4.2 Alternative Fuel Source Candidate - Hydrogen
The industry standard for aircraft fuels are different blends of kerosene with thousands of other additives like anti- corrosives, biocides, and icing inhibitors. However the use of these fuels causes a negative impact on the environment.
The global aviation industry produces 2% of all human induced carbon dioxide, while aviation produces 12% from all transport sources [22]. Furthermore, with depleting oil reserves, stringent ICARE goals, and the overall green house effect, it is imperative that alternative fuel sources be explored. One of these is a renewable energy source, compressed hydrogen. Hydrogen is a high-energy clean burning fuel whose main combustion byproduct is water vapor and trace amounts of nitrogen oxides. Furthermore, converting already existing turbofan engines is relatively simple to do with- out drastically changing the engine configuration.
Hydrogen gas has a significantly lower density than Kerosene, thus storing it in this state is not feasible due to weight and size restrictions. When compressed and cooled, the gas converts to a liquid which significantly increases the density (0.089 g/L for gas, 71 g/L for liquid) and reduces required tank size. However this liquid state must be converted into a gas for efficient combustion. This can be accomplished by installing heat exchangers at various location of the engine. These heat exchangers aim to raise the overall temperature of the flowing fuel, converting the fuel to the desired state while also reducing the specific fuel consumption(SFC). This lowered SFC in conjunction with the higher energy density of Hydrogen results in a 64.7% reduction in SFC when compared to Kerosene [24].
Hydrogen engines also run cooler than conventional kerosene engines which results in a longer life for the the engines as components wont face stringent heat wear. The final benefit of using hydrogen comes from the saved cost in the form of carbon tax, the renewable nature and relatively harmless emissions make it exempt from such levies.
Further analysis is being done on the safety implications using this fuel as well as a trade-off analysis between the volume/weight of hydrogen compared to the volume/weight of kerosene for the reference mission of 700 nm.
36 5 Weight and Balance & Component Weights
For the designed aircraft, Class I weight and balance analysis was performed by following Roskam Part II [11]. This included calculating the center of gravity locations for each component of the aircraft. In addition, the sizing of the horizontal and vertical tail as well as landing gear configuration were obtained by using the initial weight and balance analysis performed. For the initial analysis, several assumptions were used to obtain the preliminary weight break- down of the aircraft in order to locate the most aft and most forward c.g. location.
The initial weight breakdown of the aircraft is tabulated below. Notice that the weight indicated below does not correspond to the total take-off weight of 566,000 lbs, due to the omission of crew weight, as well as due to the difference between the final size and weight of the empennage compared to the initial assumptions. Furthermore, the use of hydrogen will cause an increase in structural weight in the form of larger fuel tanks, insulators, and heat exchangers, this weight was not accounted for in the initial weight breakdown therefore a contingency was included in the takeoff weight.
Table 9: Initial Weight Breakdown of SRHC-530G
Component Assumptions Weight xc.g. zc.g. Wing - 62,260 lbs 128.09 ft -2.25 ft Horizontal Stabilizer - 14,368 lbs 229.90 ft 30.00 ft Vertical Stabilizer - 19,513 lbs 211.00 ft 9.24 ft Engines - 80,000 lbs 177.47 ft 3.00 ft Fuselage Group - 56,600 lbs 119.90 ft 0.00 ft Landing Gear Same c.g. as the wing 22,640 lbs 128.09 ft -2.25 ft Baggage c.g. at the fuselage c.g. 12,300 lbs 119.90 ft 0.00 ft Passengers c.g. at the fuselage c.g. 82,000 lbs 119.90 ft 0.00 ft Fuel all fuel is stored in the wings 178,179 lbs 128.09 ft -2.25 ft Total Weight 527,860 lbs
Per Torenbeek table 8-5 [25], the fuselage was estimated to be roughly 10% of takeoff weight, while the empennage is roughly 5% of the takeoff weight, while the wings were estimated to be 11% of the take off weight. According to the figure below, the static margin of the aircraft was chosen to be 5% and then plotted against the horizontal tail sizing. This tail sizing was then used to determine the size of the vertical tail as well as individual weight contributions.
While the weight breakdown of the aircraft shown in the previous table is preliminary, the horizontal and vertical tail were sized to ensure a longitudinally stable aircraft. Therefore, the sizing process was iterated to account for the c.g. change as the weight of the empennage was adjusted to fit the desired stability limits. More detailed explanation on the sizing of empennage can be found in the stability and control section.
37 Figure 18: Required Static Margins for Different Aircraft Models [11]
Within the mission profile of the SRHC-530G, it is expected that the aircraft will undergo various loading conditions that will possibly shift the center of gravity. The loading conditions considered during the construction of the weight excursion diagram are: empty, empty plus fuel, empty plus payload, and take-off weight. The final weight excursion diagram can be seen below:
Figure 19: Weight Excursion Diagram for SRHC-530G
From the figure, it can be seen that the center of gravity shift between the most aft and most forward case is less than
0.8 ft. This is ideal as the smaller shift in center of gravity will ensure that the static margin of the aircraft will not be significantly impacted by the variance in loading conditions during the mission of the aircraft.
38 The aircraft landing gear contributes 22,640 pounds to the total aircraft weight. The placement and configuration of the landing gear heavily depends on the location of the most aft center of gravity. Along the x-axis of the aircraft, the main landing gear must be placed well aft of the xcg location to meet the longitudinal tip-over criterion. The typical angle between the main gear (on the ground) and the xcg to satisfy this criterion is at least 15 degrees. The angle used in this aircraft’s design is 23 degrees, which places the main landing gears 8.96 feet aft of the xcg and 143.98 feet aft of the nose. The nose gear is placed 110.61 feet fore of the xcg, which is 24.40 feet aft of the nose. Additionally, the lateral tip-over criterion of the aircraft must be met, which requires the lateral tip-over angle, , to be less than or equal to 55 degrees. This criterion is heavily dependent on the location of the ycg and the track of the main landing gears. The lateral tip-over angle used in this aircraft is 51.98 degrees. The outside landing gears will have a track that is 38 feet apart and there will be a third main landing gear in the center of the two outside main gears. All of the landing gears will be placed on a strut that is 12 feet high from the ground to the bottom of the fuselage.
39 6 Aerodynamics
6.1 Aerodynamics Overview
The aerodynamics design and analysis includes determining the geometrical properties of the wing that are ideal for the aircraft size and configuration as well as simulating the flow over the wings using CFD for a comprehensive lift and drag analysis. Simulations for different wing configurations and flight stages are performed to ensure optimized performance of the wing in terms of lift and drag for all flight stages. Table 10 lists the aerodynamic requirements from the initial sizing. These requirements are used in conjunction with wing configurations of existing aircraft to obtain preliminary geometrical definitions for the SRHC - 530G wings. These parameters are shown in Table 11.
Table 10: Aerodynamic Requirements
Parameter Value Cruise Lift Coefficient CL 0.51 Max Lift Coefficient Take-Off 2.14 Max Lift Coefficient Landing 2.69 Wing Surface Area [ftˆ2] 4717
Table 11: Wing Geometrical Parameters
Parameter Value Wingspan Unfolded[ft] 233.2 Wingspan Folded[ft] 210 Aspect Ratio 11.5 MAC [ft] 20.23 Sweep Angle 20 t/c 0.11 Taper Ratio 0.176 Dihedral 0 Incident Angle 2 Airfoil Boeing Airfoil J
40 The wing planform sketched in OpenVSP is shown below:
Figure 20: Wing Planform
Moreover, the aircraft will utilize variable wingtip angle technology to have different cant angles for different flight phases. The wing will be able to be folded completely for taxing, then switched to an angle of 45 for climbing, and
15 for cruise. This design choice helps drag minimization as it is realized that traditional winglets are not beneficial for all different flight phases, but mainly climbing. The airplane generates the most lift during climb, which causes significant lift-induced drag. The winglets help minimize the induced drag by interfering with the wingtip vortices.
In addition, the wingtips also increase profile drag. Hence, it is only viable to use wingtips when the induced drag minimization is greater than the profile drag addition. So it is essential for the wingtips to be optimized for different
flight phases for effective drag reduction. The drag trade-off with and without winglets can be seen in Figure 21 [32].
Therefore, with the addition of the variable wingtip angle design, the airplane will be able to reduce as much induced drag as possible during climb, while maintaining the same for cruise, which is a more critical flight phase when analyzing costs and efficiency. The optimal angles for climb and cruise were settled upon through CFD analysis and consideration from a research study[32]. An example flight plan is shown in Figure 22. Whereas this flight plan is designed for a different airplane through the research study, this plan is very similar to the expected flight plan for the SRHC-530G, which will be revised when final CFD testing is conducted. Lastly, the structural requirements for changing the wingtip angle during flight has not been considered in detail and will be discussed in the future.
41 Figure 21: Profile Drag Difference During Climb and Cruise from Traditional Winglets
Figure 22: Tentative Flight Plan for Different Wingtip Angles
42 6.2 Airfoil Selection and Characteristics
The airfoil selection criteria includes being classified as supercritical and/or a transonic airfoil, and having a thickness ratio t/c of about 0.11. The online Airfoiltools database was used to for determining airfoil candidates. The airfoils in the table below shows the selected airfoil candidates.
Table 12: Airfoil Trade Study (Re 40million) ⇡ Airfoil L/D AoA BOEING 737 OUTBOARD 91.232 3.065 BOEING AIRFOIL J 92.958 2.012 LOCKHEED C-5A BL1256 85.075 2.673 LOCKHEED C-141 BL426.57 92.467 2.838 Grumman/Gulfstream GIII Transonic 79.149 4.551 Boeing KC-135 Transonic 82.975 2.965 NASA/Langley Whitcomb Supercritical 81.560 0.485 NASA SC(2)-0010 89.793 4.272 NASA SC(2)-0410 90.559 2.123 NASA SC(2)-0610 87.876 0.501
These airfoils were evaluated for having a high lift-drag ratio during cruise as well as a low cruise alpha. These airfoils were first converted to MSES geometry format(required for XFOIL). This format requires the geometry to begin at the trailing edge and go to the leading edge over the top surface, and then back to the trailing edge through the lower surface. Then the airfoils were processed in XFOIL(ran through Python) at the cruise condition.
The cruise condition involves a cruise velocity of 0.8 mach, or 778 ft/s and the atmospheric properties at an altitude of
35,000ft. Using the average chord, the Reynold’s number was calculated to be approximately 40 million. The results for each airfoils were separated for the cruise CL, coefficient of lift for level flight, which is 0.5143. The airfoil cruise lift-drag ratio and angle of attack are also shown in the above table. The airfoil with the highest lift-drag ratio during cruise was chosen as the wing airfoil for the design. This airfoil is chosen to be Boeing Airfoil J, with a lift-drag ratio of 93 and an angle of attack of 2 at cruise. The airfoil geometry, lift curve and drag polar for cruise are shown in the following figures. This airfoil will be further analyzed in the future for transonic performance(using MSES) as a check that this is the most optimal airfoil for the SRHC.
Figure 23: Boeing Airfoil J Geometry
43 Figure 24: Lift Curve at Cruise for Boeing Airfoil J (Re 40million) ⇡
Figure 25: Drag Polar at Cruise for Boeing Airfoil J (Re 40million) ⇡
44 6.3 High-Lift Systems
In order to bridge the gap between the maximum coefficient of lift in the clean configuration and the maximum coeffi- cient of lift needed for takeoff and landing, high lift devices will be added to the aircraft. These devices, while adding the needed lift for certain flight conditions, also add weight, complexity, and cost to the aircraft and must be designed with this in mind. The two pertinent areas of the aircraft that benefit the most from high lift devices are the leading and trailing edges of the wing. This is because the simplest methods to increase the lift of the aircraft is by increasing wing area and wing camber.
Trailing-edge flaps are critical to almost every commercial aircraft in use today. They have different deployment set- tings during takeoff and landing and many planes would not be able to do either without them. Fowler flaps are the most widely-used flap design in the commercial industry due to their relative simplicity and their increase in wing area and camber. Fowler flaps extend out on tracks and often have a series of slots to add energy to the airflow to keep the flow attached over the flaps. The most common of which are the single-slotted, double-slotted, and triple-slotted
Fowler flaps. The more slots there are increases the total lift and drag produced, but also greatly increases weight, complexity, and cost. The most desirable quality of these flaps is that different flap settings are ideal for different airplane configurations. For example, at low flap deployment levels, the Fowler flaps extend from the wing and in- crease wing area and thereby lift considerably, while only creating a slight increase in drag which is ideal for a takeoff configuration. At high deployment levels, the flaps extend more in a downwards motion and create a little higher lift, but also a large increase in drag which is ideal for a landing configuration.
The trailing-edge high lift system was sized in an iterative process by choosing an initial flap to wing area ratio and comparing the resulting change in coefficient of lift to the required change in coefficient in lift for both landing and takeoff conditions. The maximum clean aircraft coefficient of lift at landing conditions was found to be 1.598 and the required aircraft coefficients of lift for takeoff and landing are 2.14 and 2.69, respectively. These values were input into the following equations to provide the change in lift based on the deflection angle of single-slotted Fowler flaps. This process was done for both landing and takeoff configurations, but for simplicity only landing is shown in the equations.
C =1.05(C C ) (1) Lmax,L Lmax,L Lmax S clmax,L = CLmax,L K⇤ (2) Swf
2 3 K⇤ =(1 0.08cos (⇤ c ))cos 4 ⇤ c (3) 4 4
cl = cl↵↵ f f (4)
45 cl determines the incremental aircraft lift coefficient available for a given flap deployment. cl↵ is the slope of the lift curve and the other two variables are found in the figures below.
Figure 26: Definition of Flapped Wing Area [12]
Figure 27: Sectional effectiveness for single-slotted flaps and Fowler flaps [40]
The process was iterated several times to optimize the high lift sizing to the desired performance and the outcome revealed a flapped wing area to wing area ratio of 0.4. The finalized aircraft coefficients of lift depending on flap deflection angle are plotted below.
46 Figure 28: Airplane Lift Curve At Landing Conditions with different Flap Deflections
It can be seen that the aircraft lift curves all somewhat follow a slope of around 2⇧. The coefficient of lift required for takeoff is achieved between ten and twenty degrees of flap deployment; the coefficient of lift required for landing is achieved at 30 degrees of flaps. It is always important to design the high lift devices for the trailing and leading edge together, as they must work together and compliment each other during different stages of flight. Single-slotted
Fowler flaps are to be selected to decrease the complexity and weight of the wing, alongside a leading-edge device to delay stall and allow for higher coefficients of lift necessary for takeoffs and landings.
Looking at the leading edge of the wing, the two most viable designs to increase lift are Krueger flaps and slats.
Krueger flaps are typically cut-out of the cruise-wing geometry and extend forwards during takeoff and landing to increase wing camber and area. The folding, bull-nosed Krueger flaps can provide a larger maximum coefficient of lift than slats can, however, they also have a smaller lift-to-drag ratio during takeoff than slats do. When deployed,
Krueger flaps create a very blunt leading edge that becomes ideal to be placed in the thicker (inboard) portions of the wing. Slats are simple leading-edge devices that, when deployed, increase the camber of the wing while keeping the leading edge of the airfoil thin. Doing so assists in lowering the stall speed and decreasing the takeoff and landing ground roll of the aircraft. The differences in deployment methods are shown in the figure below.
47 Figure 29: Diagram of Krueger Flaps and Slats [28]
Based on the description given above, the SRHC-530G aircraft will implement simple slats on the leading edge to decrease the risk of stall and decrease groundroll. Slats were chosen because they are simple, light, and cheap, as well as because the flaps have been found to provide enough coefficient of lift for the takeoff and landing requirements without implementation of leading edge devices.SLats will provide an increase in coefficient lift as a safety margin without an excessive increase in cost or weight due to the mechanical simplicity.
48 6.4 Airplane Drag Breakdown
The main components of the aircraft that contribute significant flat plate area are the wings, fuselage, horizontal tail, vertical tail and engine nacelles. Parasitic drag and wave drag calculators from OpenVSP(Appendix 11.3) were used for these calculations. The drag breakdown results shown are for cruise at Mach 0.80, and density and viscosity at
35,000 ft altitude.
Table 13: Drag Breakdown for Cruise (Open VSP)
Wetted Equivalent Percent of Item C Area Area, f (ftˆ2) D0 Parasitic Drag
Wing 8381.28 24.50 0.00519 35.68% Fuselage 11726.06 20.98 0.00445 30.56% Horizontal Tail 1047.59 3.39 0.00072 4.94% Vertical Tail 1199.98 3.50 0.00074 5.10% Engine Nacelle 1977.01 9.43 0.002 13.73% Total f, without cooling drag - 61.80 - 90% Total f, with 10% cooling drag - 67.98 - 100% Total CD0 - - 0.0131 -
49 6.5 CFD Analysis
Flightstream [41] was used to perform CFD analysis on the SRHC-530G to determine its performance by generating lift curves and drag polars. Although the chosen software is high-fidelity and quite accurate for its aerodynamics results, results inconsistencies were found which are being troubleshooted through iterative CFD analysis and redesign of the aircraft. The following figure shows the aircraft in cruise condition with separation markers to demonstrate the
flow over the aircraft during clean flight. It can be seen that there is a slight flow separation at the wing trailing edges and at the roots of the wings and tail. This could be an effect of imprecise mesh generation, so the mesh will be repaired and the analysis will be redone. The overall results of the anaylsis are discussed in this section.
Figure 30: Flow Separation during Cruise
50 6.5.1 Wingtip Cant Angle Trade Studies
First of all, the design was solidified by running a wingtip cant angle trade study to determine the angle produces the least drag. For cruise conditions, the wingtip angle was varied from 0 to 60 degrees in increments of 15 degrees and drag coefficients were obtained as shown in the table below.
Table 14: Wingtip Cant Angle Trade Study for Cruise Condition
Cruise Total Lift and Drag Coefficients Cant Angle Angle of Attack CL CDo CDi 0 1.75 0.5119 0.0239 0.0158 15 1.75 0.5087 0.0238 0.0124 30 1.85 0.5151 0.0239 0.0147 45 1.9 0.5101 0.024 0.0152 60 2.1 0.5154 0.0239 0.0159
It can be seen that for the chosen case parameters, the lowest induced drag occurs at 15 degrees. This angle was also the most optimal for Onera M6 wing from the research study as well [32]. In fact, this wingtip orientation reduces induced drag by over 25% compared to no wingtips or the traditional winglet optimized for climbing. This signifies the importance of this design choice. This wingtip angle will be chosen for cruise conditions. Furthermore, this angle could be further dialed up on generating a curve with higher resolution of different wingtip angles to determine the one that has the absolute minimum drag.
Moreover, a trade study for optimal wingtip angle for take-off and landing conditions is to be conducted as well, The initial results from the study were found to be not consistent or accurate so the study will be conducted later. For now, a wingtip angle of 45 degrees will be chosen, similar to the optimal angle of the Onera M6 [32].
51 6.5.2 Airplane (3D) Lift Curves (cruise, takeoff, landing)
Flightstream results for chosen wingtip angles and appropriate atmospheric conditions at take-off and landing and cruise are shown below. The airspeed used for the take-off and landing curve is set as 133 knots in the solver.
Figure 31: 3D Lift Curve for Take-off and Landing, Wingtip Angle = 45
Figure 32: 3D Lift Curve for Cruise, Wingtip Angle = 15
52 6.5.3 Airplane Drag Polars (Cruise, Takeoff, Landing)
With the same solve cases as the lift curves discussed above, the drag polar results are shown below. It should be noted that the drag coefficient displayed is only induced drag, not profile drag Furthermore, the induced drag is seen to decrease abruptly and reach negative values. This could be a result from erroneous solver settings. Therefore, the solver settins will be set again in order to obtain better results.
Figure 33: 3D Drag Polar for Take-off and Landing, Wingtip Angle = 45
Figure 34: 3D Drag Polar for Cruise, Wingtip Angle = 15
53 7 Stability and Control
Chapters 8 and 11 of Roskam [12] Part 2 were used to determine the static stability and control of the aircraft in a Class
1 analysis. This analysis deals with the horizontal and vertical tail sizing, as well as an aircraft trim diagram. It has been decided that the empennage will be designed with a cruciform configuration due to the rear-bodied location of the engines. The cruciform tail allows for the horizontal stabilizer to be out of the disturbed flow from the engines, while also providing undisturbed flow over both the horizontal and upper portion of the vertical stabilizer at high angles of attack to ensure aircraft stability and control.
7.1 Horizontal Tail Sizing
The determination of the horizontal tail sizing has a large impact on the longitudinal stability of the aircraft. The sizing of the horizontal stabilizer is dependent on the location of the most aft center of gravity (CG) and aerodynamic center.
The most aft CG remains relatively constant with the change in horizontal sizing, while the change in aerodynamic chord can be determined using Eq. 1 below. A horizontal tail sizing can then be assumed based off of the static margin found in Eq. 2.
CL↵h d✏h Sh X¯AC wb + (1 ) ⌘hX¯AC h CL↵wb d↵ S X¯AC = (5) CL↵h d✏h Sh 1+ (1 ) ⌘h CL↵wb d↵ S
SM = X¯ X¯ (6) AC CG
The results of equation 1 are plotted in Figure 33, along with the most aft CG values for an aircraft with a takeoff weight of 566,000 pounds. Using a conservative estimate and following Roskam’s procedure, a minimum static margin of 5% is chosen and yields a horizontal tail with a wetted area of 1047.59 ft2.
The chosen static margin from the figure above will determine the aircraft’s longitudinal static stability and will have a direct effect on the design and weight of the aircraft. The point at which the aerodynamic center and the center of gravity are the same is known as the neutral point. If the horizontal tail sizing was chosen to indicate this value, the aircraft would have neither static stability or instability and would be unreactive to perturbations in flight. Instead, transport planes are designed to have a static margin that promotes longitudinal stability such that an increase in pitch generates a nose-down pitching moment and vice-versa. A larger static margin would make the plane more longitudinally statically stable, however, a smaller static margin will decrease the weight and make it easier for the pilots to control the pitch of the aircraft. It is also important to note that the CG will move less than a foot during its most aft and most forward configurations, indicating that this static margin will be suitable for all configurations of
flight.
54 Figure 35: Horizontal Tail Sizing and Static Margin for a Takeoff Weight of 566,000 lbs
7.2 Vertical Tail Sizing
The vertical stabilizer controls the stability of the aircraft in the yaw axis, or the directional axis. A stable vertical tail is designed aft of the most aft CG location so that when encountering wind, it will inherently yaw the aircraft into the relative wind. The sizing of the vertical tail is shown in the figure below, and is based on an assumption from Torenbeek [25] such that the total empennage weight should be 5% of the total takeoff weight of the aircraft, as described in Figure 53 in the Appendix.
Figure 36: Vertical Tail Weight Sizing
Having already sized the horizontal stabilizer and finding the weight of the vertical stabilizer to be 13,450 lbs, the total
55 wetted area of the vertical tail can then be calculated to be 1199.98 ft2 via relating the horizontal tail area to its weight
[26].
7.3 Aircraft Trim Diagram
The prior empennage sizing has determined that the SRHC-530G is longitudinally stable. The next step is to determine how the aircraft must be trimmed at a cruise angle of attack to achieve the necessary coefficient of lift. This was done by calculating the aerodynamic pitching moment of the aircraft with different elevator deflection angles. First, the aircraft pitching moment slope is calculated in equation 6.3 using the lift curve slope and static margin. Then, the effect of the elevator deflection on the airplane lift coefficient and pitching moment coefficient are found using equations 6.4 and
6.5, respectively.
x x C = C ( cg ac,A )= C SM (7) m↵,A L↵,A c¯ c¯ L↵,A ⇤
Sh qh dCLh Sh qh dCLh d↵h Sh qh Sh qh CL = CLh = e = e = CL↵h⌧ e (8) S q d e S q d↵h d e S q S q 1 1 1 1
l C = C h (9) m L c¯
Important values for the equations above are found in Table 6.1 and extracted from [27]:
Table 15: Parameters for Aircraft Trim Calculations
Parameter Value 1 CL↵,A 0.1086 (deg ) 2 Sh 1047.59 ft 1 CL↵h 0.0733 (deg ) ⌧e 0.60 e -5, 0, 5 (deg)
56 Figure 37: Aircraft Trim Diagram for Cruise Conditions at W = 566,000 lbs and a CG of 0.857 of MAC
The figure above plots the aircraft pitching moment against the aircraft lift coefficient for the most aft CG condition.
As discussed in Table 3, the necessary aircraft lift coefficient for clean cruise is 0.51. This figure shows that at a trimmed elevator deflection angle of -5°, the aircraft coefficient of lift is 0.51 with an aircraft pitching moment of zero.
This means that if the aircraft is trimmed to have an elevator deflection angle of -5°, the plane will maintain steady cruise conditions without pilot input (assuming no winds). Additionally, it can be seen that the slope of the lines is
-0.05, which verifies that the slope is equal to the static margin stated prior. This negative slope is also an indication of longitudinal static stability.
57 7.4 Control Surface Sizing
The sizing of the control surfaces on the SRHC-530G were based off of similar transport jet designs. The design process was iterative in that the control surfaces were sized and the pitch, roll, and yaw tendencies were then analyzed to determine effectiveness. The control surfaces must be able to control the aircraft in adverse environments without excessive input force from the pilot.
The rudder is the control surface on the vertical stabilizer and controls the yaw of the aircraft. Due to the cruciform tail design, the rudder is divided into two sections that have independent control systems but only one set of controls for the pilot. This adds a layer of safety should one of the rudder systems become ineffective. Based on the Boeing
777-200, the ratio of the rudder area over the vertical tail are is 0.26 [36]. This would indicate that the rudder has an area of 312 ft2 based off of the 1200 ft2 area of the vertical stabilizer.
The ailerons are the control surfaces on the outboard section of the wing that control the roll of the aircraft. Typical ailerons on commercial jets, such as the B-777-200, have an aileron chord to wing chord ratio of 0.22, an inboard span ratio of 0.32, and an outboard span ratio of 0.76. This would indicate that the aileron chord is 4.45 feet, the inboard span is 34 feet, and the outboard span is 80.5 feet.
The elevators are the control surfaces on the horizontal stabilizer and control the pitch of the aircraft. It is important to note that because of the cruciform design, the elevator will have a larger aerodynamic impact due to the larger moment arm than most aircraft configurations. This means that elevator will be more effective and will need less surface area than the 777-200. Based on these factors, the ratio of the elevator area over the horizontal stabilizer area is 0.30. This provides the elevator with a surface area of 315 ft2.
The precise geometries of the SRHC-530G control surfaces still need further analysis to optimize performance and effectiveness. However, the sizes have been chosen from similar aircraft with similar missions. This should indicate that the chosen values are a strong starting point for further optimization once a higher level analysis can be completed.
58 7.5 Pitch, Roll, and Yaw Characteristics
The ability to control an aircraft is a very important factor to take into account when for control surface sizing and is often an iterative process. The pitch, roll, and yaw characteristics can also be defined as longitudinal, directional, and lateral stability and maneuverability. The FAR Part 25 requirements are to ensure the plane has sufficient control power to maintain steady cruise flight, the airplane can be safely maneuvered from one steady state flight condition to another, the cockpit control force levels are acceptable under all conditions, and the airplane can be trimmed in certain
flight configurations.
The longitudinal stability and maneuverability has already been analyzed in sections 6.1 and 6.3 and will not be dis- cussed in further detail in this section besides the fact that it has proven to pass FAR Part 25 requirements. The lateral and directional maneuverability is dependent on aileron and rudder areas and maximum deflection angles and will be further analyzed following Part VII of Roskam for Determination of Control, Stability, and Performance Character- istics[40]. The process will be iterative and the control surface sizing and maximum deflection angles will change to optimize performance. The first iteration of control surface sizing can be found in section 6.4, and the maximum deflection angles based off of similar aircraft with similar missions can be found in the table below.
Table 16: Maximum Control Surface Deflection Angles
Control Surface Max. Deflection Angle (deg) Elevator Up 25 Elevator Down 30 Aileron Up 30 Aileron Down 10 Rudder 27.3 ±
59 8 Performance
8.1 Takeoff and Landing
8.1.1 Takeoff
The RFP requires the aircraft to takeoff within a maximum of 9000ft field length. This distance is calculated by sum- ming the acceleration, rotation, transition and climb-out distances with a 15% buffer[34]. The table below summarizes the results. The take-off distance is found to be 7433.16 ft, so the distance is within the limits.
Table 17: Takeoff Performance Analysis Results
VS VLOF SA SR STR SCL STO 389.95 428.94 5424.3 857.89 20.99 163.18 7433.16
8.1.2 Landing
For landing, the RFP requires a maximum distance of 9000ft similar to takeoff. The landing distance is calculated by summing the Approach, free-roll, and braking distances with a 40% buffer [35]. The main results are shown below.
The landing distance was calculated to be 8597.7 ft, which is acceptable under the requirement.
Table 18: Landing Performance Analysis Results
VS VA VTD SA SFR SB SL 183.85 239.00 211.42 2203.8 634.27 2320.4 8597.7
60 8.2 Climb Requirements
Figure 38: Thrust Required Vs Thrust Available
A thrust required curve was plotted against the available thrust generated by the twin GE9x setup. As further specifi- cations of these engines become readily available (i.e Brake Horse Power & Propulsive Efficiency), the rate of climb of the aircraft can be accurately computed. However, using the max excess thrust and the minimum required thrust,
ft the max rate of climb is then found to be 200.41 min .
61 8.3 Cruise Performance & Payload-Range
The cruise performance of the aircraft can be defined by analyzing the payload-range diagram for the airplane [42].
The following payload-range diagram was obtained using the TSFC of GE90 engines and L/D values at cruise obtained through OpenVSP, assuming constant lift coefficient and constant mach number for the cruise stage:
Figure 39: Payload-Range Diagram for SRHC-530G
It can be seen that the range of the aircraft with maximum payload is 3787 nmi, which includes 200 nmi extra range for flight to an alternate airport and 1 hour loiter. When the fuel is maximized by 125% of the fuel used during normal operation, the range goes up to 5636 nmi. The assumption for the maximum fuel capacity is preliminary for this stage and will be justified through structural limits in future work. When the aircraft does not carry payload and has maximum fuel, the range is determined to be 7128 nmi.
62 8.4 Comparison with Baseline/Competing Aircraft
The designed aircraft can be compared to the selected baseline aircraft that was used in the initial sizing of the aircraft.
It is important to note that the values presented in the table below will be fine tuned in the next stage of the design process. Below, SRHC-530G is compared to competing aircraft: Boeing 777-200 and Boeing 777-300ER.
Table 19: Comparison of SRHC-530G with Competing Aircraft [6]
Specification Boeing 777-200 Boeing 777-300ER SRHC-530G Flight Crew 2 2 2 Capacity 400 (2-class) 451 (2-class) 414 (2-class) Length 209.1 ft 242.3 ft 239.8 ft Wingspan 199.9 ft 212.6 ft 233.2 ft Wing Sweep Back 31.6 31.6 24.0 Wing Area 4605 ft2 4702 ft2 4717 ft2 Cabin Width 19.3 ft 19.5 ft 18.2 ft Fuselage Width 20.3 ft 20.3 ft 19.3 ft Empty Operating Weight 297,300 lbs 370,000 lbs 296,000 lbs Maximum Take-off Weight 545,000 lbs 775,000 lbs 566,000 lbs Cruise Speed Ma = 0.84 Ma = 0.84 Ma = 0.80 Maximum Range 5,240 nmi 7,930 nmi 3,787 nmi Takeoff Distance at MSL 8,300 ft 10,500 ft 7,450 ft
8.5 Comparison with Baseline/Competing Aircraft
The aircraft design is also compared to the given mission specs from the RFP.
Table 20: Comparison of SRHC-530G with Mission Specs
Specification Mission Specs SRHC-530G Flight Crew 2 2 Capacity 400 (2-class) 414 (2-class) Maximum Approach Speed 145 KCAS 140 KCAS Design Range 3,500 nmi 3,787 nmi Takeoff Distance at MSL 9,000 ft 7,450 ft Landing Distance at MSL 9,000 ft 8,598 ft
In addition, the aircraft carries 5 cubic feet per passenger for baggage and meets the galley, lavatory, and exit require- ments provided in 14 CFR Part 25. The design range also incorporates the required reserve fuel specifications.
63 9 Structural Layout, Materials, Manufacturing
9.1 Aircraft Structure, Material, and Manufacturing Choices
The SRHC-530G will use the most advanced available materials and manufacturing technologies in the aerospace industry. This design will differ from comparable aircraft by primarily using composite materials rather than con- ventional metals. Sections of the aircraft that will be using conventional metals will have the most progressive forms of that material. The materials will be manufactured in a process that incorporates the most sophisticated industrial technologies that are economical and have the greatest rate of production.
In many comparable airplanes, aluminum is the material of choice due to its mechanical properties, low density, and low cost. However, many of the parts once occupied by aluminum are being replaced by composites due to its lighter weight, improved fatigue strength, and fuel efficiency. Composites have a high stiffness, strength, and corrosion resis- tance, whereas aluminum is corrosion sensitive and high strength alloys are not wieldable. Despite the advantages of composites, the cost of raw materials and manufacturing is expensive, as well as maintenance due to delaminations.
Nevertheless, the long term performance, weight reduction, and decrease of parts makes composite use the ideal ma- terial for the aircraft’s major structures.
The design will use a combination of carbon sandwich, carbon laminate, and fiberglass composites. Major structural sections of the aircraft will use carbon laminate composites, such as the fuselage, wing box, and stabilizer boxes.
Secondary structures will use carbon sandwich composites such as the rudder, elevators, wingtips, and nacelle cowl- ing [43]. Fiberglass epoxy laminates will be used for the wing fairings, stabilizers, radome, and the wing-to-fuselage fairings. The aircraft interiors will use graphite composites in the floor panels, main deck side walls, ceiling panels, and overhead stowage bins.
In order to minimize the financial impact of using composite materials, the manufacturing process will use the most innovative technologies available that have a high production rate, low cure time, and low waste factor. The best in- dustrial processes for aerospace composites are the prepreg layup and liquid composite molding. The prepreg layup is further broken down into hand layup and machine layup processes. The hand layup process is responsible for produc- ing fairings, vertical stabilizer skin panels, and wing control surfaces. The machine layup process is responsible for producing fuselage barrels, fuselage skins, stabilizer skins, wing skins, wing spars, and structural stiffeners. This will be accomplished by using Automated Tape Lamination, fiber placement, and drape forming technology. The liquid composite molding process will contribute to parts like ribs, rods, bars, and beam sections. This will be accomplished
64 Figure 40: Aircraft Material Selection using Resin Transfer Molding, and pultrusion.
Conventional metals will be used in locations where its mechanical and material properties reign superior to composite materials for that application. Aluminum will be used in the leading edges of the horizontal and vertical stabilizers.
It will also be used on the front part of the engine cowling and leading edge of the wing. Since these locations are more vulnerable to impacts from debris, aluminum was chosen to avoid expensive composite repairs. Additionally, aluminum has better heat transfer than composites, which makes the material better to use in the wing leading edge for the anti-icing system. The most advanced aluminum alloys will be used in these locations using precipitation hardening. The alloys that will be used in the aircraft are 7140-T76511, 7136-T76511, 2098-T8, and 2099-T83 [44].
Additionally, titanium and steel will also be used in applications that need to use its high mechanical strength prop- erties. Titanium will be used in locations when the operating temperature is much greater than 270oF. Titanium has applications in the wheel wells, galleys, lavatories, fire walls, and the tail cone. This aircraft will use the titanium alloy
Ti-10V02F3-3Al. Steel will be used when very high strength is needed (275-300 ksi) such as in the flap tracks and in the landing gear system with titanium.
65 9.2 V-n Diagram
A V-n diagram can be constructed to observe the load limits of the designed aircraft for a variety of flight conditions
[45]. In the figure below, a composite diagram is shown that includes the V-n and gust diagrams for FAR 25 certifica- tion. The aerodynamic parameters used are for cruise conditions at 35,000 ft and Ma 0.8 speed.
Figure 41: Composite V-n & Gust Diagram
The airspeeds denoted in the diagram is in KEAS, equivalent airspeed. It can be seen that the gust limits, shown as dashed lines, are within the V-n diagram and the design speed at maximum gust intensity, VB, is calculated to be 150.9 KEAS, which is less than the designated cruise speed of 256.7 KEAS. This diagram is then used for determining the structural layout and material selection of the aircraft. Additional data extracted from the V-n diagram is tabulated below:
Table 21: Composite Diagram Data
Structural & Aerodynamic Parameters Positive load limit 2.5 Negative load limit -1.25 Design Maneuver Speed, VA 109.7 KEAS Design Speed at Max. Gust Intensity, VB 89.3 KEAS Design Cruise Speed VC 256.7 KEAS Design Dive Speed VD 320.8 KEAS
+1g Design Stall Speed VS1 69.4 KEAS
-1g Design Stall Speed VS 1 82.6 KEAS
66 10 Systems
10.1 Cabin Layout
The cabin layout of SRHC-530G was determined after sizing the fuselage with consideration to number of passengers, galleys, lavatories, emergency exit and passenger door width, and instructions on fuselage sizing in Roskam Part II
[11] and III [12]. The mission requirements listed below were used in the sizing:
Table 22: Cabin Layout Specifications
Aisle Width 20” Economy Pitch 32” Economy Width 18” Business Pitch 36” Business Width 27” Economy Seats 360 passengers Business Seats 54 passengers
In addition, the dimensions for the galleys and lavatories were obtained from Roskam Part III while emergency exits were sized and selected in accordance with 14 CFR 25.807 and 25.813 [4]. The chosen parameters for the galleys, lavatories, and emergency exits are as follows:
Table 23: Galley, Lavatory, and Emergency Exit Specifications
Galleys and Lavatories Galley Dimensions 79”x25” # of Galleys 5 Lavatory Dimensions 40”x40” # of Lavatories 8 Emergency Exits Emergency Exit Type Type I Emergency Exit Dimensions 24”x48” Minimum Required Cross-Aisle Width 20” # of Emergency Exits 8
Including the structural considerations as listed in Roskam Part III, the geometry of the fuselage was determined.
Note that the cockpit length is based on B707 dimensions in Roskam Part III for cockpit dimensioning. The resulting specifications for the fuselage is listed below along with the cross-sectional view of the economy and business class areas:
67 Figure 42: Economy Class Fuselage Cross-Section
Figure 43: Business Class Fuselage Cross-Section
68 Table 24: Fuselage Dimensions
Cockpit Length 11.48 ft Cabin Length 163.17 ft Fuselage Cone Angle 16 deg Fuselage Cone Length 65.14 ft Fuselage Total Length 239.79 ft Fuselage Inner Diameter 18.23 ft Fuselage Outer Diameter 19.13 ft Fuselage Outer Height 15.88 ft Structural Depth 5.38 in
The type of cargo containers that are to be used for passenger baggage is LD3-45 Unit Loading Device (ULD) [46], whose specifications are tabulated below:
Table 25: ULD Specifications
ULD Type LD3-45 # of ULDs 12 ULD Gross Weight, per 915.1 lbs ULD Volume, per 131 ft3
The cabin layout of the aircraft as a top view is shown below:
Figure 44: Cabin Layout
69 10.2 Cockpit
The cockpit of SRHC-530G is planned to integrate Honeywell Airplane Information Management System (AIMS)
[47], which is an integrated modular avionics system that has also been used in Boeing 777.
Figure 45: Honeywell AIMS [47]
Honeywell lists this product as a mission control element. It is claimed to integrate and provide cockpit display system, flight management system, thrust management system, aircraft condition monitoring system, and many more.
The maintenance period for the system is every 15 years [47]. This technology is currently in use in many Boeing
777s, and is expected to be available for SRHC-530G.
70 10.3 Emergency Egress, Fire Protection Safety Systems
The Fire Protection System for this aircraft uses separation, isolation, detection, and control to prevent and extinguish
fires from the passenger cabin, cargo compartments, engine, and APU. Different material selections and thermal in- stallations are used to minimize the spread of fires. The design of the aircraft will separate fire essentials such as fuel, ignition source, and oxygen. The design will also isolate potential fires from disseminating to other locations of the aircraft and control the fire if one occurs.
The passenger compartment of the aircraft will be made of materials that are self-extinguishing. This includes the interior ceiling, sidewall panels, partitions, passenger seat material, and electrical wires. Thermal insulation will be installed behind the cabin interior panels in order to delay fire from entering the cabin. The airplane skin and insula- tion blankets will be capable of resisting burn-through from fuel-fed exposure on the bottom half of the fuselage for a minimum of four minutes to allow evacuation of passengers before burn-through may occur. Photoelectric-Area type smoke detectors, which detects the presence of smoke particles in the air by the reflection of scattered light, will be used in the crew rest and lavatories and mounted on the ceiling or upper sidewalls. Additionally, six handheld/1211
Halon fire extinguishers will be placed throughout the cabin [48].
The freight area of the aircraft is considered an FAA Class C cargo compartment. There must be an approved built- in fire extinguishing system that is controllable from the flight deck. The design of the compartment will exclude hazardous quantities of smoke, flames, and suppression agents from entering the passenger cabin while being able to control ventilation and drafts within the compartment. The compartment design will have liners that are separate but attached to the aircraft structure to prevent the fire from spreading to other locations of the aircraft. All wires, tubing, equipment, and controls will be designed away from the compartment to minimize the fire hazard. All heat sources will be shielded and insulated to prevent ignition. The cargo compartment smoke detector is based on photoelectric sensing. The cargo compartment will use a Halon 1301 suppressant [49].
It is crucial for the engine and the Auxiliary Power Unit (APU) to have a fire suppression system integrated into its design. Both the engine and APU will have a passive and active fire suppression system incorporated in the design. The passive system includes features such as non-combustible materials, electrical grounding, use of fire walls, ventilation, drainage, and separation of fluid-carrying lines and wires. The active system will have a fire and overheat detection system, fire extinguishing system, temperature system, and automatic shutdown system. The engine will have a
flammable fluid drainage zone were flammable fluids will drain overboard using holes, hoses, and tubing. Ventilation will be incorporated to minimize the accumulation of flammable air and exit without being re-ingested. Electrical
71 components will be explosion proof and will not overheat. All electrical systems will be grounded in case of static electricity or a lightning strike. The fire suppressant system used in the engine will be Halon 1301. The APU will be isolated in the tale cone with the fire extinguishing system located in the aft fuselage. The fire-extinguishing systems for the engine and APU will be controlled from the flight deck [50].
72 10.4 Landing Gear
The landing gear system for this aircraft will consist of one strut near the nose of the aircraft and three struts near the wings. The nose gear will have two tires on the strut at 190 psi and the main gear will have four tires on each strut at
200 psi. All of the tires will be cylindrical with a diameter of 52 inches and a thickness of 20.5 inches. The main gear tires have a static load of 54,536.45 pounds. The nose gear tires have a static load of 28,897.69 pounds and a dynamic load of 48,308.78 pounds. These loads are provided to allow for greater tire customization since tire manufacturers use these loads to rate their tires. Since the aircraft will be cruising above 150 knots, the landing gear will be retractable and in a tricycle configuration to avoid a significant drag penalty. As mentioned in the Weight and Balance section, the landing gear is designed around the most aft center of gravity in order to fulfill the lateral and longitudinal tip-over criterion. The nose gear is placed 24.40 feet aft of the nose along the center of the aircraft, which is a great location for ease of maneuverability during aircraft ground operations. The main landing gear is placed 142.98 feet aft of the nose with the two outside gears having a track length of 38 feet and the third main gear being placed along the x-axis of the aircraft. In order to avoid a tail strike, the main landing gears (from the ground) must have an angle greater than
15 degrees from the most aft point of the fuselage. The current design has a ground clearance angle of 16.22 degrees, satisfying the criterion. The current placement of the landing gear is designed for the nose gear to take on 7.5% of the maximum aircraft takeoff weight and for the three main gears to take on 92.5% of the load, which meets the typical loads observed in transport jets.
73 Figure 46: Landing Gear Configuration
74 11 Detailed Design and Analysis: Hydrogen Storage
As discussed in the propulsion section, two different fuel options are considered to be implemented for the SRHC-
530G. The kerosene version was designed to completion as far as the requirements from AIAA and EAE 130 go, and its design is to be modified to work with hydrogen. In the modification, minimal changes are to be made to the original design and most things are kept the same. This approach is taken to minimize costs for the goal of slowly transitioning into long-range commercial transport using hydrogen fuel for the sake of a sustainable future for aviation. Moreover, this detailed design study focuses solely on the storage of hydrogen. However, the effects of the modifications to implement hydrogen on the original design are important to consider and are briefly discussed below.
It is realized later in Section 10.2 that hydrogen is stored in the liquid hydrogen state and at cryogenic temperatures.
Due to the physical properties of cryogenic liquid hydrogen and its highly explosive nature, it is ideal for it to be stored in the fuselage, rather than the wings. This requirement causes one major design change — location of fuel tanks. In order to minimize further major changes, the tanks are simply placed in the front and the rear of the fuselage with the goal of preserving the existing seating layout within the fuselage.
Furthermore, there are few other changes caused by the above modification which are noteworthy. The wings are not held down by fuel or engines during flight so they require extra structural reinforcements. This increases empty operating weight in comparison to the original kerosene design. However, hydrogen fuel weighs much less than kerosene so the maximum takeoff weight actually decreases. These effects are only to name a few and several other changes are expected, which will be considered in the future.
75 11.1 Sizing and Structures
The hydrogen tanks were sized for the AIAA range requirement of 3500nmi + 200nmi for loiter. This requires
8393.8ft3 of hydrogen fuel. As there are two fuel tanks to be implemented within the airplane, each fuel tank is sized
3 a to hold 4196.9ft of hydrogen. According to Gomez [56], elliptic domes with a major to minor axis ratio b =1.66 are the best compromise of weight and tank length. This conclusion is used in the sizing along with the diametral restriction from the fuselage cross section. The following numbers for tank length and diameter are obtained using the above relations:
Table 26: Tank and Fuselage Dimensions
Tank Length 39.19 ft Tank Diameter 14.73 ft Fuselage Diameter 18 ft Fuselage Height 15 ft Fuselage Length 240 ft
The sizing numbers conclude that there is a few inches of vertical clearance and few feet of horizontal clearance within the tank and fuselage. This should be sufficient for the structures within the fuselage where the tanks will be placed and will require mounting. Furthermore, the tanks take up about 80 ft along the length of the fuselage, which is one-third of the fuselage length. Therefore, it is realized a range of 3500 nmi and 400 passenger capacity cannot be accomplished at the same time, and one must be compromised on. For the satisfaction of both requirements, significant redesign of the whole airplane would be required, which is not the goal of this detailed design study, so this shortcoming will be neglected for now.
11.1.1 Shell
The figure below shoes design of the tank shell developed in Solidworks. It can be seen that the tank is cylindrical rather than elliptic, which is done to minimize manufacturing costs. The optimized a/b ratio of 1.66 is still applied.
Furthermore, the shell is 2mm thick and reinforced by a frame as discussed below. The material of the shell is discussed in a later section as well.
11.1.2 Frame
SRHC-530G hydrogen tanks are designed to withstand the pressure required for cryogenic storage. As cryogenic storage is chosen over high-pressure storage, tanks can be sized as low-pressure vessels. In addition to the skin, the tanks include a two-layer frame structure that consists of outer rings and inner longitudinal stringers. The tank skin is attached to this frame. The frame structure modeled on SolidWorks 2019 [63] can be seen below:
76 Figure 47: Hydrogen Tank Shell - Iso View
Figure 48: Hydrogen Tank Frame - Iso View
Table 27: Hydrogen Tank Frame Dimensions
Parameter Value Frame Length 39.19 ft Frame Diameter 14.73 ft # of Outer Rings 24 # of Longitudinal Stringers 36 Stringer Thickness 1.0 in Stringer Width 1.0 in
The frame matches the shape of the tank skin and has two domes and a cylindrical section. There are 24 outer rings and 36 longitudinal stringers. This initial frame design is iterative and depending on the FEA results, the structure can be altered to make it stronger in the case of failure.
The tank frame can be seen in top and side view as well:
77 Figure 49: Hydrogen Tank Frame - Side View
Figure 50: Hydrogen Tank Frame - Top View
78 11.2 Storage
The main problem with storing hydrogen on aircraft is its extremely low density compared to normal jet fuel. Storing a sufficient amount of hydrogen calls for large storage tanks that increase the weight and decrease the usability of an aircraft. So, in order to make hydrogen propulsion practical, the storage method must be able to increase the density of hydrogen to decrease the size and weight of the storage tanks.
The two main methods of hydrogen storage are high pressure storage and cryogenic storage. The high pressure method stores hydrogen as a gas at pressures up to 10,000 psia and is the most conventional and widely-used storage option.
As storage pressure increases, the density of hydrogen also increases and this pattern is shown in the figure below.
Figure 51: Hydrogen Density at Various Pressures and Temperatures [58]
It is important to note that as pressure increases, the storage tank thickness must be increased to withstand the applied stresses. It also needs expensive and heavy systems equipment to maintain the tanks at high enough pressures to hold enough hydrogen for the 3500 nautical mile mission requirement. This, along with the fact that high pressure storage is quite dangerous and could lead to loss of life, are the reasons cryogenic storage has been chosen for the SRHC-
530GH.
79 The cryogenic storage system allows for a reduction in tank weight and volume by storing the hydrogen as a liquid.
Liquid hydrogen has a density about 70 times larger than gaseous hydrogen at standard atmospheric pressure, but it must be kept below the boiling temperature of Hydrogen of 21 Kelvin. To further increase the density above that of liquid hydrogen, the hydrogen can be stored as a half-and-half mixture of solid and liquid hydrogen, called slush hydrogen, which has a density 80 times larger than gaseous hydrogen [59]. The effect that the temperature and pressure of the hydrogen is stored at is shown below, indicating there is no need for extreme pressurization of the tanks.
Figure 52: Cryogenic Temperature, Pressure, and Density Values [59]
Cryogenic storage maximizes the amount of hydrogen to be stored in the smallest possible volume and weight, how- ever, it does not come without its difficulties. Because it must be maintained at cryogenic temperatures, an airtight insulation system must be used to reduce the boil-off rate. Cryogenic hydrogen also has a limited lifetime due to the boil-off rate and will most likely have to be produced on-site. The fuel tanks must also be constantly sealed and pres- surized (around 25 psia) to reduce boil-off rates and ensure no air or liquid gets inside which would instantly freeze and clog fuel lines. That being said, the only two significant systems involved in storing the cryogenic hydrogen are the tank and insulation [61].
Recent research in cryogenic storage has also shown potential for laminar flow control by use of the sensible heat of the stored fuel to cool aerodynamic surfaces. The cryogenic temperatures allow for a cooling of the boundary layer on the wings and nacelles to delay the transition to turbulent flow and reduce the frictional drag. Although the implementation of surface cooling increases the aircraft operating empty weight through addition of heat exchangers, pumps, and control subsystems, the laminar flow control has shown to be a feasible and cost-saving application of cryogenic hydrogen storage [62].
80 11.3 Materials and Insulation
11.3.1 Materials
In order to safely carry liquid and gas hydrogen, the tank needs to be made of materials that have high strength, high fracture toughness, high stiffness, low density, and low permeation to liquid and gas hydrogen [59]. A trade study was performed between aluminum, steel, titanium, and composites to determine the best material for the tank. The material and mechanical properties are listed in Table 27.
The material selection process considers the properties listed in Table 28 as well as the cost of the material and its manufacturability. 304L steel was identified as a potential material; however, its material properties were undesirable due to its high density and low yield strength compared to other materials. Also, steel performance is unreliable in cold temperatures, which is a concern because the hydrogen needs to be stored in cool tanks. TC4 titanium was also analyzed and its material properties were in a desirable range with a low density, high yield strength, and low ther- mal conductivity. However, the raw material and manufacturing cost of titanium is unfavorable. Composite materials showed favorable properties such as in the density, yield strength, and thermal conductivity. Despite the potential for a significant reduction in weight, the costs of production and development are high for composite materials compared to conventional metals [58]. Also, there are potential safety concerns about permeation with hydrogen and being able to identify fractures [59].
After identifying all the potential materials, the 209X series aluminum was selected as the choice material. The
209X series advanced aluminum is an ideal material because of its low density and its yield strength is in a desirable range. Aluminum is also reliable in cold temperatures. Although aluminum has a high thermal conductivity, it will be insulated with a composite material to keep the tanks cool. Aluminum is also affordable to acquire and manufacture.
For the FEA analysis, the aluminum’s elastic modulus was entered as 78 GPa and a Poisson’s Ratio of 0.34.
Table 28: Properties of Select Materials[64]
Material Density Yield Strength Thermal Conductivity 3 Aluminum (209X) 2.63 g/cm 470 MPa 164 W (m K) 1 3 Steel (304L) 7.93 g/cm 410 MPa 15 W (m K) 1 3 Titanium (TC4) 4.5 g/cm 825 MPa 7.96 W (m K) 1 3 Composite 1.53 g/cm 1900 MPa 0.00675 W (m K) 1
11.3.2 Insulation
LH2 needs to be stored at low temperatures to keep the chemical composition as a saturated liquid/gas mixture.
However, during various stages of flight, the tank is exposed to numerous heat sources from its surroundings, causing
81 variations in tank pressure and temperature. This in turn causes LH2 boil off rates to increase. Acceptable boil off rates for aircraft are in the order of 0.1% by weight per hour or less [59]. Furthermore, to minimize the operating empty weight of the aircraft, low density foams such as aerogels and multi-layer insulation (MLI) are the most desirable.
MLI consists of alternating layers of low conductivity spacer and low emissivity foil, and offer thermal performance an order of magnitude better than foam. Unfortunately, this performance degrades rapidly for pressures higher than
100 mPa, posing an unnecessary safety risk. The proposed tank uses polyurethane foam for insulation material.
Figure 53: Fuselage and Tank Layers
The insulation can be applied to the interior or exterior of the tank wall. If the insulation is placed outside the tank, diffusion of gaseous hydrogen will prevail and prevent effectiveness. Furthermore the insulation is also more suscep- tible to mechanical damage as the tank volume increases and decreases at various stages of flight. Lastly, since the liquid hydrogen is at cryogenic temperatures, direct exposure to the tanks can cause some undesirable effects to the structural integrity of the tank. Thus interior insulation is chosen as shown above.
Figure 54: Placement of Insulation and Liners
Aluminium is used as the primary material to compose the storage tank, this stems from the fact that Aluminium shows minimal susceptibility for hydrogen embrittlement. The interior insulation being used is inner wetted thermal
82 insulation spray, a spray on polyurethane foam with embeded metallic liners, specifically designed for this use case.
This foam has a thermal conductivity of 0.00675 W/m K [56], acceptable for our use case.
Lastly, as mentioned before variations in altitude and pressure result in a net energy gain/loss to the storage tanks, along with the heat gained from the catwalk of the storage tanks, the tank storage pressure is subject to fluctuate. To maintain storage pressure, a venting pressure of 25 psi is used. Thus pressure was selected as it provided the most aerodynamic and cost effective option, a lower venting pressure would result in the loss of significant amounts of hydrogen fuel while higher venting pressures requires thicker structures.
11.4 Structural Analysis
Structural analysis was performed on the tank frame to determine the feasibility of its design under pressure loads. The pressure loads are due to the cryogenic storage of hydrogen at 22 psi. To be conservative, a pressure load of 25 psi was applied on the outer rings of the frame where the tank skin is attached. Also, the structure is fixed at the domes seen in green in Figure 55 to reduce external connections to the tank that would compromise the structural integrity. Using the specified aluminum discussed in materials section, a finite element analysis was performed on SolidWorks FEA
[63]. Only a pressure load is applied on the system whereas the low temperature effect on the material is discussed in the previous Materials section. Prior to the SolidWorks analysis, an attempt was made to run the tank frame FEA on
PATRAN/NASTRAN, however, there were issues with model meshing that could not be solved. In the figure below, the frame deformation plot can be seen along with the stress heat map of the structure.
Figure 55: Tank Frame FEA Stress Heat Map and Deformation Plot
83 Table 29: Hydrogen Tank FEA Results
Parameter Value Maximum Displacement 0.0215” Maximum Stress 5.66 ksi Al 209X Yield Strength 68.2 ksi
The maximum displacement of the frame under 25 psi pressure load acting on the outer rings is 0.0215 inches while the maximum stress is 5.66 ksi. Compared to the material yield strength of 68.2 ksi, the structure is expected to withstand the pressure load without plastic deformation. Therefore, this tank frame design is final. However, future work could be done on the loading cycle of this structure to determine the number of cycles it can endure to determine the life of the structure.
11.5 Conclusions and Future Work
In the previous pages, a detailed methodology for implementing and designing the storage system for hydrogen propul- sion is outlined and described in detail. The completed work has been extrapolated from in-depth research of certain portions of the liquid hydrogen system to ensure that our design is using cutting-edge technology while still making an impact on cost and emissions. The primary hurdle with hydrogen storage comes from the loss of fuel through gaseous diffusion, this happens when various heat sources are introduced to the tank causing the LH2 to boil - off. Furthermore, contact of the cold LH2 against the skin of the tank can cause some undesirable material effects. To prevent this from happening an interior insulation made of polyurethane foam is used.
The tank was designed to hold enough fuel for a little over 3700 nautical miles per the AIAA requirement, as well as to cryogenically store that fuel as liquid hydrogen. The hydrogen tank structure was tested for failure using SolidWorks
FEA, and results indicated that the structure was safe for low pressure, cryogenic storage. In addition to the stress due to pressure loads, the structure was also equipped with insulation to protect it from temperature loads. In the future, the tanks can be evaluated for product life to determine maintenance schedules and make sure that it is a clean and safe technology to replace kerosene in aviation.
84 12 Cost and Utilization
12.1 Capital Costs
Research, development, testing and evaluation (RTD&E) along with production cost was computed using the method- ologies detailed by Raymer [51]. Per the RFP, an assumption of 15 planes per month for five years was assumed to compute these values. All equations used are listed in the appendix.
Using the empty aircraft weight, the maximum aircraft speed,and the total number for production, the airframe en- gineering, tooling, manufacturing, and quality control hours were computed. An assumption of 5 flight test planes were then used to determine the flight test cost. Lastly, the development support costs, manufacturing materials costs, engine and avionics costs were then used to determine the total RTD&E as well as flyaway cost was computed for
five years of production. Values obtained were then adjusted for inflation in 2020. The assumptions for various hourly costs are summarized below:
Labor Wage ($) Engineering Hourly Cost 115 Tooling Hourly Cost 118 Hourly Manufacturing Cost 98 Hourly Quality Control Cost 108
Table 30: Labor Costs
The table below details the sale price of the SRHC - 530G for the first five years of production, the plot on the next page summarizes how RTD&E, as well flyaway cost decreases as the number of units produced increases:
Year Sale Price ($ Million, 15% profit added) RTD&E + Flyaway Cost ($ Million) 1 242 211 2 210 183 3 185 161 4 170 148 5 170 148
Table 31: Annual Acquisition and Development Price in millions ($)
85 Figure 56: Cost Per Unit Adjusted for Inflation with 15% Profit
86 12.2 Utility Cost
To accurately represent the cost of the aircraft, the direct operating cost should also be considered in addition to the
flyaway cost of SRHC-530G. Direct operating cost of the aircraft includes fuel oil, maintenance, crew, depreciation, rentals, insurance, and other related costs. Fuel oil, maintenance, and crew costs are considered as variable direct operating costs while the rest is considered to be fixed. To estimate the direct operating cost of SRHC-530G, the cost report published by FAA [52] is used. In the table below, the average direct operating cost breakdown for widebody aircraft with more than 300 seats is shown.
Table 32: Average Direct Operating Cost per Block Hour for Widebody Aircraft Adjusted for 2020 US Dollars [52]
Category Cost per Block Hour Fuel and Oil $11,517 Maintenance $1,891 Crew $1,724 Total Variable Cost $15,132 Depreciation $853 Rentals $356 Insurance $10 Other $6 Total Fixed Cost $1,225 Total $16,357
The values indicated are for air carriers that file Schedule P-5.2Carriers Filing Schedule P-5.2 is considered for air carriers with total operating revenues of $100 million or more per year. In addition, the direct operating cost break- down can be shown graphically. Below, the pie charts are included for variable, fixed, and total direct operating cost breakdown.
87 Figure 57: Variable Direct Operating Cost per Block Hour
Figure 58: Fixed Direct Operating Cost per Block Hour
88 Figure 59: Total Direct Operating Cost per Block Hour
The graphs indicate that the major portion of the direct operating costs is fuel oil. Thus, SRHC-530G aims to reduce the cost and impact of fuel by using alternative energy sources, such as hydrogen.
The designed aircraft is suited for short range (700 nm) as well as medium range (3500 nm) flight. Therefore, a direct operating cost analysis is performed for the two reference missions with the indicated range.
Table 33: Direct Operating Cost analysis for 700 nm and 2500 nm reference missions
Reference Mission 700 nm 3500 nm Average Speed Ma 0.8 Ma 0.8 Flight Time 1.62 hours 8.10 hours Total Block Hours 1.97 hours 8.45 hours Total Direct Operating Cost $32,210 $138,149 Cost per Available Seat Mile $0.1111 $0.0953
The table also includes the cost per available seat mile for each reference mission. This is an important parameter for aircraft carriers as it indicates the operational cost of an air carrier to carry one seat a distance of one mile. It can be seen that the cost per available seat mile is lower for a longer range mission due to the lesser impact of additional ground and air operations on a longer flight time.
The annual direct operating cost of the aircraft can be calculated by estimating the block hours of the aircraft in a year.
This value is estimated using the daily block hour average provided by MIT Global Airline Industry Program - Airline
Data Project [53], that is 12.44 hours for widebody aircraft average for American air carriers. The resuls are tabulated below.
89 Table 34: Annual Direct Operating Cost Analysis
Daily Block Hours Yearly Block Hours DOC per Block Hour Yearly DOC 12.44 hours 4540.6 hours $16,357 $74,270,594
12.3 Comparisons to Competing Aircraft
Aircraft commonly used to service the short to medium range family were compared with the acquisition and operating costs of the SRHC-530G. This sector is commonly serviced by the Airbus A320-200, Boeing 737 - 800, Boeing 757 -
200, Airbus A330-200, and Boeing 777 - 300. The unit price for each is:
Aircraft Seats Unit Price ($, Millions) Airbus A320-200 143 101 Airbus A330-200 250 238 Boeing 737 - 900 220 109 Boeing 757 - 200 189 100 Boeing 777 - 300 304 307
Table 35: Unit Cost of comparable Aircraft [55]
From a unit cost perspective the cost of the SRHC - 530G fares well against the competition listed above. Furthermore, most of the aircraft listed above have seating capacities vastly lower than that of the RFP, yet the unit cost of 170 million (after 4 years of production) competes well provided it has more seating, thus more efficient. While finding direct operating costs for each of the above aircraft was difficult, the utility cost analysis done in section 10.2 outline the major operational costs. The major operational costs come from fuel and the fact the SRHC can take hydrogen which potentially eradicates carbon tax paid by the airliner as well as meeting various emission requirements.
12.4 Lifecycle CO2 Analysis
Aviation is responsible for roughly 12% of all CO2 emissions from transport sources and 2% of all human induced
CO2 emissions [22]. This figure is expected to increase as the global aviation industry grows. Furthermore CO2 emissions are present in the manufacture, assembly, and recycling of such aircraft. Finding adequate manufacturing data for the various components of the aircraft is not readily available, furthermore emissions data for the GE9x isn’t either as the engine isn’t in commercial use it. However while conducting the trade study between hydrogen and kerosene it was found that roughly 4000 gallons of kerosene would service this trip. Data from the US Energy
Information Administration cites that roughly 22 pounds of CO2 is released per gallon of kerosene combusted. Thus, assuming perfect combustion, 88,000 lbs of CO2 would be released on a 700 nmi flight. However since the SRHC can accommodate hydrogen, using hydrogen fuel can be provide a CO2 emission free flight as the byproduct of hydrogen combustion is water [54].
90 13 Conclusions and Recommendations
SRHC-530G by AVIAVIS is a short range, high capacity aircraft that can carry 414 passengers in a two-class layout.
Its range is optimized for 700 nautical miles in terms of cost in the design stages. Currently, it is also capable of traveling a distance of 3700 nautical miles with maximum payload, with even more increased range in maximized fuel loading conditions. For the propulsive systems, a rear mounted twin engine configuration was chosen with the newly developed General Electric GE9x engines. 19% of an airline’s operating cost comes from fuel; therefore, to meet the RFP requirements of a cost-efficient aircraft an engine with a high bypass and pressure ratio was selected.
Furthermore, to reduce the overall carbon emissions and save capital on carbon tax, alternative fuels are also being considered. A trade off study was conducted on the viable use of hydrogen as a propellant, for a reference mission of
700 nmi it was found that roughly 11,900 gallons of liquid hydrogen would provide the energy required to transport a fully loaded aircraft. However, in order to make this feasible, certain key technologies need to come into effect.
For efficient storage, fuel tanks located in the wings need to be able to withstand the high pressures to ensure that the hydrogen remains in liquid state. Heat exchangers need to be developed with the specialized purpose of heating the hydrogen without causing an explosion. Lastly infrastructure in airports as well as hydrogen production plants need to be scaled, improved, and/or developed to meet the strong demand poised by the airline industry.
A tentative wing configuration was established with key features being a wing area of 4717 ft2 and a variable winglet cant angle setup. The wingtips will be at different angles during different flight missions for optimized aerodynamic performance. The cruise wingtip angle was chosen to be 15 through a CFD trade study. Additional Flow analysis was performed for different configurations and lift curves and drag polars were obtained. However, for the takeoff and landing curves, flaps and landing gear were not included in the simulation, but only the varying wingtip angle, velocity, and air properties were considered. Proper analyses with flaps, wingtip angles, and landing gear will be performed using Flightstream [41] in the future. Furthermore, the landing gear in its current configuration meets the lateral and longitudinal tip-over criterion. The landing gear is also placed to meet the ground elevation criterion and provide easy ground maneuverability for operators. To finalize the design procedure for the landing gear systems, research will be conducted on the retraction capability and its integration within the aircraft fuselage. The SRHC-530G’s primary and secondary structures will be made of composite materials. Conventional metals will be used in applications where the mechanical and material properties are superior to composites. Composites are expensive in their raw form and to manufacture; therefore, the most sophisticated industrial technologies will be used for efficient production.
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96 15 Appendix
15.1 Fuselage Calculations
Unit Loading Device
Number of ULDs = V (n + n /V ) baggage ⇤ crew passenger ULD
WeightofeachULD = W (n + n ))/n + W baggage ⇤ crew passenger ULD ULDempty
Economy Class
Total Width of Economy Class = width n economy ⇤ abreasteconomy
Number of Economy Rows = neconomy/nabreasteconomy
Length of Economy Rows = n pitch roweconomy ⇤ economy
Business Class
TotalWidthofBusinessClass = width n business ⇤ abreastbusiness
Number of Business Rows = nbusiness/nabreastbusiness
Length of Business Rows = n pitch rowbusiness ⇤ business
Width of Fuselage
Cabin Width = width + width n economy aisle ⇤ aisle
Fuselage Inner Diameter = (width )2 +(h )2 2 cabin seat ⇤ p
97 FuselageStructuralDepth =0.02 d 12 + 1inches, fromRoskamP artII ⇤ i ⇤
Fuselage Outer Diameter = d + depth 2 i structural ⇤
Length of Cabin + Cockpit
Ltotal = Lcockpit + Lgalley + Llavatory + Lbusiness + Leconomy + Lemergencyexit
Length of Fuselage Cone
1 FuselageConeLength = do/tan (✓)
Length of Aircraft
Laircraft = Lcabin+cockpit + Lcone
Fineness Ratio
Rfine = Laircraft/do
98 15.2 Weight Excursion Diagram Calculations
Empty Weight: W1
W1 = Wempty
Empty Weight + Payload Weight: W2
W2 = Wempty + Wpayload
Empty Weight + Fuel Weight: W3
W3 = Wempty + Wfuel
Take-off Weight: W4
W4 = Wempty + Wfuel + Wpayload
Then, center of gravity location for each loading condition is calculated and the resulting C.G. is plotted along with the aircraft weight.
99 15.3 Payload-Range Diagram Calculations
The payload-range diagram was constructed by assuming constant lift coefficient and constant Mach number. Thus, the following equations were used for different loading conditions:
a CL W0 Rcruise = M ln( ) c CD W1
Point A
R =0
Wpayload = Maximum P ayload
Point B
a CL W0 RB = M ln( ) c CD W1
W0 = MTOW
W1 = OEW + Maximum P ayload + Reserve F uel
Wpayload = Maximum P ayload
Point C
a CL W0 RC = M ln( ) c CD W1
W0 = MTOW
W = MTOW Maximum F uel + Reserve F uel 1
100 W = MTOW Maximum F uel OEW payload
Point D
a CL W0 RD = M ln( ) c CD W1
Wpayload =0
Figure 60: Payload Range Diagram Example
101 15.4 Drag Breakdown: Parasitic Drag
Figure 61: OpenVSP Parasite Drag Calculation
102 15.5 Aircraft Part Sizing Based on Max Takeoff Weight
Figure 62: Aircraft Sizing as Percentage of MTOW [21]
103 15.6 Calculating Various Landing Gear Parameters
The location of the xcg is measured from the nose to 127.36 ft. into the fuselage. The ycg is measured from the coordinate axis of the ovular fuselage up 2.59 ft. The variable bf , is measured from the middle of the fuselage to the bottom of the fuselage. The value 0.65 is the distance between the middle of the fuselage to the coordinate axis. The height of the landing gear may be selected by the engineer, but may need to change to meet tip-over criteria. The known variables allow for the main landing gear distance from the c.g. to be calculated (see Eqn.1). A number greater than or equal to 15 degrees will automatically satisfy the longitudinal tip-over criterion; however, some of the variables in the equation may need to be adjusted to meet other criterion.
o o lm =(ycg + hLG + bf +0.65) tan (23 15 ) [1]
Once the maximum aircraft takeoff weight and the load of nose gear are known, the value calculated from equation 1 may be placed into equation 2, which gives the fore distance of the nose landing gear to the most aft center of gravity.
lm(Pn WTO) ln = [2] Pn
The next criterion to check for is the lateral tip-over criterion. When the distance of the Track, T, is known, then the angle between the main landing gear to the nose gear may be calculated using Equation 3, ✓ (see Figure 52).
Figure 63: Lateral Tip-Over Criterion (Top View)
104 lm + ln ✓ = arctan T [3] 2 ! Next, the surface altitude, X, may be found since the Track and ✓ is known using Equation 4.
X = Tsin✓[4]
Using simple geometry, the length of the surface altitude of the main gear with its intersection at the x-axis may be calculated using Equation 5.
2 2 T w = (lm) + [5] 2 s ✓ ◆ Once w is found, the lateral tip-over criterion angle, , may be found using Equation 6 (see Figure 53). This angle must be less than or equal to 55 in order to meet the criterion.
Figure 64: Lateral Tip-Over Criterion (with Inclined Angle)
ycg + hLG + bf +0.65 = arctan T [6] 2 ! The next part of the landing gear configuration analysis is to determine if the location of the main landing gears satisfy the ground clearance angle criterion. In order for this criterion to be fulfilled the angle between the most aft part of the fuselage and the main landing gear (at the ground) must be greater than 15 degrees (see Figure 54). The distance between the main landing gear to the most aft location of the fuselage may be found in Equation 7.
105 mgtt = 239.79 xcg lm [7]
The longitudinal ground clearance angle may be determined using Equation 8.
Fh + hLG ⌘ = arctan [8] mgtt ✓ ◆
Figure 65: Ground Clearance Angle
106