Environautics EN 1
Response to the 2009 2010 AIAA Foundation Undergraduate Team
Aircraft Design Competition
Presented by Virginia Polytechnic Institute and State University
Left to Right: Justin Cox, Julien Fenouil, Jason Henn, Ryan Hofmeister, Michael Caporellie, Justin Camm, August Sarrol, Richie Mohan
Environautics Team Roster
ii Executive Summary
Environautics presents the EN 1 as a solution to the 2009 2010 American Institute of
Aeronautics and Astronautics (AIAA) Undergraduate Aircraft Design Competition Request For
Proposal (RFP). The design will serve as an environmentally friendly and efficient strut braced wing commercial transport to replace the Boeing 737 and Airbus 320. The RFP calls for a medium range, biofuel capable transport aircraft capable of carrying 175 passengers and cargo over a range of up to 3500 nautical miles and entering service by the year 2020. The main drivers for the proposal include maximizing performance capabilities with respect to the given RFP mission and maintaining a competitive commercial advantage while reducing the aircraft’s overall environmental impact through improved efficiency and usage of biofuels. The requirements of the RFP are discussed in Section 2.1.
The proposed design incorporates the strut braced wing design, a design proven in lightweight general aviation aircraft that enables a reduction of the weight of the main wing spar, allowing for efficiency enhancements through reduced wing thickness and sweep angle. The inclusion of advanced biofuels in the design minimizes performance penalties while reducing the aircraft’s environmental impact, further enhancing the competitive capability of the design compared to existing aircraft in areas such as operating costs. Operating costs are reduced further through use of advanced technologies that permit the aircraft to operate at increased efficiency and fewer delays.
The combination of performance, efficiency, advanced technologies and reduced environmental impact make the Environautics EN 1 a first rate choice for future commercial transports.
iii
Table of Contents Executive Summary ...... iii Index of Figures ...... viii Index of Tables ...... ix Nomenclature ...... x 1. Introduction ...... 1 2. Request for Proposal ...... 1 3. Fuels ...... 3 3.1 Fuel Descriptions ...... 3 Fuel Types ...... 3 3.2 Biological Fuel Sources ...... 8 3.3 Sizing Requirements ...... 8 3.4 Fuel Decision ...... 10 4. Concepts ...... 13 4.1 Blended Wing Body (BWB) ...... 13 4.2 Hybrid Blended Conventional ...... 14 4.3 Strut Braced Wing ...... 15 4.4 Design Selection ...... 16 5. Sizing ...... 17 5.1 Initial Weight ...... 17 5.2 Thrust to Weight and Wing Loading ...... 20 6. Aerodynamic Performance ...... 22 6.1 Drag Polar ...... 22 6.2 Lift to Drag Ratio ...... 23 6.3 Airfoil Selection ...... 24 6.4 Airfoil Analysis ...... 27 7. Propulsion ...... 32 7.1 Engine Technologies ...... 32 7.2 Engine Selection ...... 35 7.3 Fuel System ...... 35 7.4 Engine Maintenance ...... 36 8. Performance ...... 37 8.1 Takeoff, Landing, Balanced Field Length Analysis ...... 37 8.2 Mission Profile ...... 39 9. Weights and Structures ...... 40 9.1 Final Weight ...... 40 9.2 Center of Gravity ...... 43 9.3 Materials ...... 44 9.4 V n Diagram ...... 47 9.5 Structural Analysis ...... 48
vi 10. Stability and Control ...... 56 10.1 Longitudinal Stability Analysis ...... 56 10.1.1 JKayVLM Analysis ...... 57 10.1.2 Tornado Analysis ...... 58 10.2 Control Types ...... 60 10.3 Cruise Trim ...... 60 11. Systems ...... 61 11.1 Cabin Layout ...... 61 11.2 In flight Systems ...... 62 11.3 Cockpit Systems ...... 66 11.4 Ground systems ...... 67 11.5 New Advanced Systems ...... 68 11.5.1 NextGen ...... 68 11.5.2 Lidar/Optical sensing Interface ...... 69 11.5.3 GPS Landing...... 69 12. Cost Estimation ...... 70 12.1 Cost of Research, Development, Testing and Evaluation ...... 71 12.2 Cost of Manufacturing – Lifetime and Unit Cost ...... 72 12.3 Direct Operating Cost ...... 74 13. Concluding Remarks ...... 79 12. References ...... 81
vii Index of Figures
Figure 4.1 Blended Wing Body Concept ...... 14 Figure 4.2 Hybrid Blended Conventional ...... 15 Figure 4.3 Strut Braced Wing Concept ...... 16 Figure 5.2.1 T/W vs. W/S ...... 21 Figure 6.1.1 Cruise Drag Polar ...... 23 Figure 6.3.1 C l vs. Sweep Angle for various values of (t/c) and M dd ...... 25 Figure 6.3.2 Main Wing Airfoil Sections ...... 27 Figure 6.4.1 Limits of laminar flow control technologies ...... 29 Figure 6.4.2 Boeing 737 Root Airfoil Boundary Layer ...... 30 Figure 6.4.3 Boeing 737 Midspan Airfoil Boundary Layer...... 31 Figure 6.4.4 NASA SC(2) 0710 Airfoil Boundary Layer ...... 31 Figure 7.3.1 Airfoil Fuel Tank and C.G. Location ...... 36 Figure 7.3.2 Wing Fuel Tank and C.G. Location ...... 36 Figure 7.4.1 Nacelle panels open for maintenance and engine removal ...... 37 Figure 8.2.1 Mission Profile ...... 39 Figure 9.1.1 Component Weight Comparison ...... 42 Figure 9.2.1 Weight C.G. Excursion Diagram ...... 44 Figure 9.3.1 Aircraft Materials’ Cost per Pound ...... 45 Figure 9.3.2 Aircraft Materials’ Relative Density ...... 45 Figure 9.3.3 Aircraft Materials’ Relative Yield Stress ...... 46 Figure 9.3.4 Materials Used In EN 1 Body ...... 47 Figure 9.4.1 V n Diagram for maneuver and gust ...... 48 Figure 9.5.1 Layout of the wing, strut, jury strut, and vertical offset ...... 49 Figure 9.5.2 Overall layout of the strut and wing design ...... 50 Figure 9.5.3 Distributed load with empty wing fuel tanks ...... 51 Figure 9.5.4 Distributed load with full wing fuel tanks ...... 51 Figure 9.5.5 Node placement and structural components ...... 53 Figure 9.5.6 Degrees of freedom for the strut braced wing ...... 53 Figure 9.5.7 Structural layout for the EN 1 ...... 55 Figure 10.1.1 Aircraft Layout ...... 59 Figure 10.2.1 Main wing ailerons, flaps and leading edge slats ...... 60 Figure 11.1.1 Cabin Layout ...... 62 Figure 11.3.1 Cockpit Systems Layout ...... 66
viii Index of Tables
Table 2.1 RFP General Requirements ...... 3 Table 3.3.1 Fuel Data ...... 9 Table 3.3.2 Necessary Boeing 737 800 Data ...... 9 Table 3.3.3 Fuel Sizing Calculation Results ...... 10 Table 3.4 Fuel Trade Study Results ...... 12 Table 4.4.1 Aircraft Decision Matrix ...... 17 Table 5.1.1 Constants from Boeing 737 800 analysis in Nicolai’s Program...... 18 Table 5.1.2 Variables for the different aircraft designs ...... 19 Table 5.1.3 Constants for Nicolai’s sizing program ...... 19 Table 5.1.4 Weights that are determined by using Nicolai’s Program ...... 20 Table 5.2.1 FAR Requirements ...... 21 Table 6.1.1 Cruise Component Drag ...... 22 Table 7.1.1 Pratt and Whitney PW1000G specifications ...... 32 Table 7.1.2 General Electric 1854 specifications ...... 33 GENX Table 7.1.3 Rolls Royce Trent 1000 specifications ...... 34 Table 7.1.4 CFM Leap X specifications ...... 34 Table 8.1.1 Takeoff, Landing, and Balanced Field Length Results ...... 38 Table 8.2.1 Mission Segment Analysis...... 40 Table 9.1.1 Weight Statement ...... 43 Table 9.5.1 Cantilever and strut braced wing comparison ...... 56 Table 10.1.1 Numeric input parameters for the JkayVLM.exe program ...... 57 Table 10.1.2 Output from JkayVLM.exe ...... 57 Table 10.1.3 Stability Derivatives from Tornado Analysis ...... 59 Table 11.2.1a List of the systems on board for the Strut Braced Aircraft ...... 63 Table 11.2.1b List of the systems on board for the Strut Braced Aircraft ...... 64 Table 11.2.1c List of the systems on board for the Strut Braced Aircraft ...... 65 Table 12.1.1 Cost of Research, Development, Testing and Evaluation ...... 72 Table 12.2.1 Total manufacturing cost of the strut braced airplane design ...... 73 Table 12.3.1 Direct Operating Cost of Flight ...... 75 Table 12.3.2 Direct Operating Cost of Maintenance ...... 76 Table 12.3.3 Direct Operating Cost of Depreciation per Nautical Mile ...... 77 Table 12.3.4 Direct Operating Cost of Landing Fees and Registry Taxes ...... 77 Table 12.3.5 Direct Operating Cost of Financing the Airplane ...... 78 Table 12.3.6 Total Direct Operating Cost ...... 78 Table 13.1 RFP Compliance Summary ...... 80
ix Nomenclature
AR Aspect Ratio b Wing Span Cfe Skin Friction Coefficient CD0 Zero Lift Drag Coefficient CLα Lift Coefficient Slope CLmax Max Lift Coefficient CLTO Take Off Lift Coefficient d Fuselage Diameter e Oswald Efficiency Factor g Acceleration Due to Gravity G Climb Gradient Lift to Drag Ratio Q Dynamic Pressure QR Specific Energy Sexposed Exposed Surface Area Sref Wing Reference Area Swetted Wetted Surface Area Thrust to Weight Ratio Vstall Stall Speed VTO Takeoff Speed W1 Takeoff Weight W2 Landing Weight Wing Loading η0 Overall Efficiency Λ Sweep Angle to Max Thickness ρ Density σ Density Ratio
x 1. Introduction
In 2006, Boeing stated their belief that there will be a higher annual increase in passenger air travel – roughly 5% per year – compared to the annual increases in fuel efficiency, making the claim that more fuel would be needed regardless of efficiency advances [1]. Two years later, commercial transport and shipping flights in the United States surpassed 18 billion gallons of jet fuel used annually by airlines. This equates to almost 450 million barrels of crude oil at roughly
100 dollars each, or 45 trillion dollars [2]. The problems of fuel efficiency, production and pollution are more applicable than ever in today’s world. These issues must be resolved in order to maintain the future growth of the world economy.
Along with a demand for alternative fuel capabilities in aircraft, there are requirements stipulating that future aircraft have increased range, higher lift to drag ratio, and an emphasis on sustainability with a rate of fuel consumption equivalent or less than what is seen in the present industry. With these demands, there is a requirement for a new development and implementation of a lighter, more fuel efficient, practical aircraft. The need for a revolution within the airline industry regarding how it impacts society as a whole must be addressed with a cost efficient, environmentally friendly aircraft.
2. Request for Proposal
The Request for Proposal (RFP) of the American Institute of Aeronautics and
Astronautics (AIAA) calls for the design of a commercial aircraft for service in the year 2020.
This design is considered to be a replacement for the Boeing 737NG and Airbus A320 aircraft.
Improvements will be focused on integrating new technologies and alternative fuels in agreement with the National Aeronautics Research and Development Challenges, Goals and Objective [3].
1 This request for design is taking momentum from the need for alternative fuels to create environmentally friendly aircraft. The current methods of replacing petroleum based fuels through direct addition of alternative fuels gives rise to the need for advanced technologies to improve energy efficiencies. Specifications of the alternative fuels will be studied through environmental emissions, noise, and carbon footprint from the aircraft design [3]. The future cost of incorporating these fuels is also a factor that is to be coupled for ensuing ground and service support systems and airport infrastructure [3].
Requested enhancements of design from the RFP state that the aircraft will have a 25% increase in the lift to drag ratio from current aircraft being replaced. Airfoil technology will incorporate laminar flow techniques for improved transition delays on swept wings. Also desired are improved weight fractions and engine efficiency [3]. The structure is expected to be similar to the light weight composites and materials utilized in the Boeing 787 aircraft with specifications taken from manufacturer projections [3]. The general requirements, as stated by the RFP, are presented in Table 2.1.
The engines given for reference in the RFP include CFM International’s CFM56 5 or
Pratt and Whitney’s PW6122A. Engine selection will be based upon improvements projected
from engine manufacturers in accordance with the targeted alternative fuel from biofuels.
The intention of the AIAA’s RFP is geared toward improving aircraft and engine
efficiency for the future use of alternative fuels. Projections are to reduce aviation environmental
impacts of emissions, noise, and fuel burn [3]. The design will be subject to comparison with
current technologies of aircraft such as the Boeing 737NG and Airbus A320 families.
2 Table 2.1 RFP General Requirements [3]
Design Factor Requirement Safety and Airworthiness Regulations FAR 25 Crew 2 Passengers 175 (1 Class) Seating Pitch 32”, Width 17.2” Width > 12.5 ft Cabin Dimensions Height > 7.25 ft Cargo Volume 1,240 ft 3 Takeoff Distance 8,200 ft Landing Speed < 140 KCAS Maximum zero fuel weight plus fuel Maximum Landing Weight reserves for maximum range Cruise: Mach 0.8 Operating Speed Maximum: Mach 0.83/340 KCAS Initial: 35,000 ft Cruise Altitude Maximum: 41,000ft Nominal: 1,200 nm Range Maximum: 3,500 nm Payload Capability 37,000 lb
3. Fuels
3.1 Fuel Descriptions
A critical aspect of the RFP is the requirement that alternative fuels are used for the design of the aircraft. These alternative fuels should be more environmentally friendly than standard Jet A 1. The alternative fuel used must be able to perform efficiently in all flight conditions the aircraft will encounter. When choosing a fuel, the type of fuel and the fuel source must be considered.
Fuel Types
There are a wide range of fuel types to consider. The primary focus of alternative fuel studies for the near term typically centers on biofuels. This report will discuss the advantages and
3 disadvantages of traditional Jet A 1, Fischer Tropsch synthetic fuel, liquid hydrogen, liquid methane, methanol, ethanol, butanol, biodiesel, and synthetic paraffinic kerosene.
Jet A-1
Jet A 1 is the current standard for aviation fuel. Any proposed replacements have to come close to meeting the standards that Jet A 1 has set. Jet A 1 has a relatively high energy density and a low freezing point suitable for the cold temperatures encountered at high altitudes. Current engines are designed to use Jet A 1 and thus current and future engines may need to be redesigned for an alternative fuel. Current Jet A 1 has a major disadvantage in the current aviation industry in that it is a fossil fuel. The fuel is derived from crude oil and is thus highly polluting, particularly in regard to carbon dioxide emissions. In the current industry climate, environmental concerns have taken a larger role in design decisions and have hastened the search for an alternative [4].
Fischer-Tropsch Synthetic Fuel
Fischer Tropsch (F T) synthetic fuel is based on the concept of the Fischer Tropsch process to convert a synthetic gas into liquid hydrocarbons. The fuel produced can have similar properties to Jet A 1, and research has shown that the energy density of the fuel is slightly higher than Jet A 1. F T synfuel is considered a drop in fuel because it can either be blended with Jet A
1 or used on its own with little to no adverse effects on fuel and engine performance. This has made it a leading contender for a short term replacement for Jet A 1. However, the fuel is considerably more environmentally destructive since it is still a fossil fuel derived from coal.
While emissions directly from the aircraft engine are slightly lower than Jet A 1, the process used to create the fuel is more polluting than the process to create Jet A 1, resulting in a net
4 increase in released greenhouse gas emissions of 147%. The price to produce F T synfuel is comparable to current crude oil prices between $80 and $100 per barrel. [8][30]
Liquid Hydrogen
Liquid hydrogen is one of the main contenders for long term aviation fuel solutions. The fuel has a pollution value of nearly zero, and has a large specific energy, but the energy density is the lowest of all the alternative fuels. This means that hydrogen fuel will have a lower mass than any other fuel but require the largest volume to contain it. Liquid hydrogen needs to be cryogenically stored, meaning the tanks need more insulation, which increases the weight and volume. The fuel would also require a completely redesigned infrastructure to accommodate the fuel [4].
Liquid Methane
Liquid methane has many of the same advantages and disadvantages as liquid hydrogen.
In addition, liquid methane, while plentiful around the world in ice deposits on the ocean floor, is difficult and dangerous to acquire and transport [4].
Methanol
Methanol is an alcohol fuel meaning it is mildly corrosive to the current infrastructure used to store aviation fuel. It is, however, a partial drop in fuel because it can be blended with aviation fuel and only requires small modifications to the engine. Methanol has a low specific energy and low energy density compared to other alternatives, which makes it less desirable as a fuel.
5 Ethanol
Ethanol is another alcohol based fuel. It is a popular alternative fuel suggestion in many fields, particularly within the automotive industry. The increased interest in production for other industries would decrease the time it would take for economic viability. In addition, ethanol has been researched extensively as a fuel. While most of these studies are done for automobiles, some of the results are still applicable to aviation propulsion systems. It has less of an environmental impact when burned compared to most fossil fuels, particularly if it is biologically derived. It is also suited for blending with some other fuels to improve performance and minimize problems with the fuel properties. This is useful as a stop gap until a fuel ultimately replaces traditional Jet A 1 [9]. Ethanol has a moderately low specific energy and energy density resulting in a large increase in the fuel’s mass and volume required. Ethanol on its own also has problems with its clouding point. The fuel will start to gel well before the minimum operating temperature standard set by Jet A 1. For ethanol to be a viable alternative, it must be further developed to overcome these problems [4].
Butanol
The third alcohol based fuel is n butanol, referred to as butanol throughout this report.
Butanol’s corrosiveness is significantly less than ethanol, making it the most attractive of the alcohol fuels for storage and transport. The process to create butanol is similar to ethanol, which is attractive as the processes for ethanol production are well documented and widely used. Like ethanol, it can be blended with many other fuels to increase efficiency and help mitigate problems with fuel properties. Butanol’s specific energy and energy density values are typically much higher than ethanol and come close to the values for biodiesel. In addition, as a stand alone fuel it does not have as many issues with freezing and clouding that ethanol and biodiesel have.
6 It also resists contamination by water better than other alcohol based fuels. As an aviation fuel, butanol has potential if it is researched further [10].
Biodiesel
Biodiesel is one of the most favored alternative aviation fuels because of its properties. It has high specific energy and energy density values. Aviation fuels can be mixed with biodiesel, or biodiesel can be used by itself, without a significant performance penalty. The fuel is not corrosive to engine parts like some alcohol based fuels. Biodiesel has a number of disadvantages, however. The foremost is its current inability to withstand the temperatures encountered at high altitudes without clouding or freezing. Blending with Jet A 1 can reduce or eliminate this problem, but this is not a long term solution for a Jet A 1 replacement. It also has a high flash point compared to Jet A 1, which means it is harder to ignite the fuel [9][11].
Synthetic Paraffinic Kerosene
Synthetic paraffinic kerosene (SPK) is a newly developed fuel alternative that can be produced from biological sources. The production process uses the Fischer Tropsch method as discussed earlier, but avoids the pollution caused by coal. This fuel is chemically the same as Jet
A 1 and is not corrosive, preventing the need to change the current infrastructure. Also, the specific energy and energy density are comparable to Jet A 1. SPK has passed all initial certification tests that Jet A 1 must undergo. SPK has been in several 50% by volume flight tests with Jet A 1, and there was no significant change in engine performance. Currently, SPK is capable of using many different biomass sources, allowing the use of the best source available.
More research is necessary to develop a commercial production facility, but BioJet Corporation has recently been under agreement with a major distributor to sell 4 million barrels over the course of 2 years [25].
7 3.2 Biological Fuel Sources
The benefit of using biological fuel sources to reduce the environmental impact of an airliner centers around the consumption of greenhouse gases during the growth of the biomass.
Greenhouse gases are still released during operation, but the consumption of the same gases during growth results in an essentially carbon neutral process. These fuel sources can generally be broken down into three groups for biofuels: first generation, second generation, and third generation biofuels. First generation biofuels are typically made from feedstock such as corn, soybeans and the seeds and grains of wheat. They are widely criticized for diverting world food supplies, particularly in poor areas of the world. These sources also have a low amount of energy produced per acre of land. Soybeans can produce about 60 gallons of biofuel per acre of land.
Second generation biofuels are produced from non food biomass, such as waste biomass from crops and cellulose. However, like first generation fuels, the land required to grow the biomass is high for the amount of fuel produced [11]. Third generation fuel sources require less land to produce the same amount of fuel compared to the first two generations. Currently, algae derived
fuel is the only source to be placed in this category. Predictions suggest that algae could produce
150 to 300 times the fuel that an equivalently sized crop of soybeans could produce in the same
timeframe. Algae can be grown on land not suitable for most other crops, so it can avoid using
land that would be used for producing food or other essential products. In addition, algae are
capable of producing a wide variety of different fuels. These advantages have made it a leading
contender for a future alternative fuel source [4].
3.3 Sizing Requirements
An important consideration in meeting the RFP is the use of biofuels and how they affect the sizing requirements for an aircraft. Using 8 common biofuels and Jet A 1, it was possible to
8 determine the requirements for an aircraft meeting the RFP through the use of the Breguet range equation and knowledge of a Boeing 737 800’s capabilities. Table 3.1 provides information on each fuel studied. The Breguet range equation is
(3.1) Range = ln The information presented in Table 3.2 represents the necessary information to determine the
overall efficiency times the lift to drag ratio. The landing weight was determined by subtracting
the fuel weight, except for the reserve fuel, from the takeoff gross weight. By using Jet A 1 fuel,
it was found that η 0 was equal to 2.8 for a Boeing 737 800. Assuming the improved of 25%
specified by the RFP, η 0 was multiplied by 1.25 to obtain 3.48 for an aircraft capable of meeting the RFP lift to drag ratio improvements.
Table 3.3.1 Fuel Data [10][12][30]
Specific Density Energy Density Life Cycle Emissions Fuel Type Energy (Hp 59°F (Hp hr/ft^3) (% change from Jet A 1) hr/lb) (lb/ft^3) Jet A 1 7.3 50.4 368 0 FT Synfuel 7.5 47.4 354 147 Liquid Hydrogen 20.3 4.4 90 6.5 Liquid Methane 8.4 26.5 224 NA Methanol 3.4 49.7 167 NA Ethanol 4.6 49.6 228 23 n Butanol (Bio) 6.2 50.6 313 NA Biodiesel (typical) 6.6 54.3 357 68 Jatropha/Algae SPK 7.47 46.7 349 109
Table 3.3.2 Necessary Boeing 737 800 Data [13][14]
Fuel Weight (lbs) 47,000 Fuel Volume (gallons) 6,900 Takeoff Weight (lbs) 174,000 Landing Weight (lbs) 127,000
9 By using the specific energies in Table 3.1, the new η 0 value of 3.48, and holding the landing weight constant, is was possible to determine the weight of fuel required to fly the 3500 nm maximum range specified by the RFP. The range was adjusted to 4000 nm to allow for the required 45 minute extra flight time required by FAR Part 25 when flying at night. From the weight of the fuel and the density in Table 3.1, the volume required to hold the fuel was also calculated. A summary of these results is displayed in Table 3.3.
Table 3.3.3 Fuel Sizing Calculation Results
% Mass Mass Volume % Volume Fuel Type Change from (lbs) (Gal) Change from 737 737 Jet A 1 52,000 8,000 9.8 9.8 FT synfuel 50,000 8,000 6.8 13.7 Liquid Hydrogen 16,000 27,000 66.0 287.2 Liquid Methane 43,000 12,000 8.1 75.1 Methanol 150,000 23,000 219.4 224.2 Ethanol 95,000 14,000 101.5 105.0 n Butanol (Bio) 64,000 9,000 35.4 35.1 Biodiesel (typical) 59,000 8,000 25.3 16.3 Jatropha/Algae SPK 50,000 8,000 6.8 15.3
3.4 Fuel Decision
The chart used to perform the fuel selection is given in Table 3.4. The metrics were chosen based on known information about the properties of each fuel and their importance to the decision process. The trade study uses a weighted scale to calculate the relative desirability of each fuel. It is a 10 point scale intended to give a good indication of how each fuel fares relative to the others. The weights were chosen based on how important each metric is to meeting the proposal’s design requirements. The most highly weighted metrics were the volume required, mass required, and environmental impact for each fuel.
10 Each fuel was given a score based on its ability to meet the requirements, where a higher value is better. The values given were relative to Jet A 1, which would score a ‘5’ for each of the metrics and give a baseline for comparison. If, for a particular fuel, a metric was not directly applicable, the fuel would score a ‘5’ to prevent skewing the results. The weight values were used as a percent out of 100 to produce a final score that is easily comparable to the baseline value. The scores were multiplied by the weight to get a weighted score, and then each of the scores for a fuel were added together to calculate the final score. Based on the results of the trade study, SPK was the clear choice to use as the fuel for this design. It scored well above all of the other fuel choices, since it is similar to Jet A 1, and scored better than the baseline due to its reduced environmental impact.
11 Table 3.4 Fuel Trade Study Results
Criteria Weight (%) FT Synfuel Liquid Hydrogen Liquid Methane Methanol Points Weighted Points Weighted Points Weighted Points Weighted Drop In Capability 4 5 0.2 0 0 0 0 2 0.08 Blend Capability (with Jet A 1) 6 5 0.3 0 0 0 0 3 0.18 Volume Required Compared to Jet A 1 20 5 1 1 0.2 3 0.6 1 0.2 Mass Required Compared to Jet A 1 15 6 0.9 8 1.2 7 1.05 1 0.15 Ease of Ignition 5 5 0.25 8 0.4 8 0.4 4 0.2 Cold Weather Capability (Unblended) 9 5 0.45 8 0.72 8 0.72 3 0.27 Cold Weather Capability (Blended) 5 5 0.25 5 0.25 5 0.25 4 0.2 Infrastructure Redesign 7 5 0.35 1 0.07 1 0.07 3 0.21 Ease of Acquiring/Production 6 4 0.24 1 0.06 2 0.12 4 0.24 Overall Environmental Impact 15 0 0 9 1.35 6 0.9 6 0.9 Safety Hazards 8 5 0.4 1 0.08 1 0.08 4 0.32 Totals 100 4.34 4.33 4.19 2.95
Criteria Weight (%) Ethanol Butanol Biodiesel Jatropha/Algae SPK Points Weighted Points Weighted Points Weighted Points Weighted Drop In Capability 4 3 0.12 4 0.16 5 0.2 5 0.2 Blend Capability (with Jet A 1) 6 5 0.3 5 0.3 5 0.3 5 0.3 Volume Required Compared to Jet A 1 20 2 0.4 4 0.8 5 1 5 1 Mass Required Compared to Jet A 1 15 2 0.3 4 0.6 4 0.6 6 0.9 Ease of Ignition 5 6 0.3 6 0.3 2 0.1 5 0.25 Cold Weather Capability (Unblended) 9 3 0.27 4 0.36 3 0.27 5 0.45 Cold Weather Capability (Blended) 5 4 0.2 5 0.25 4 0.2 5 0.25 Infrastructure Redesign 7 3 0.21 4 0.28 5 0.35 5 0.35 Ease of Acquiring/Production 6 5 0.3 4 0.24 4 0.24 4 0.24 Overall Environmental Impact 15 6 0.9 7 1.05 6 0.9 8 1.2 Safety Hazards 8 4 0.32 5 0.4 5 0.4 5 0.4 Totals 100 3.62 4.74 4.56 5.54
12 4. Concepts
4.1 Blended Wing Body (BWB)
The blended wing body concept was developed primarily to take advantage of the decreased drag due to the blended surfaces reducing the overall surface area, the increased lift due to the airfoil shaped fuselage, and the reduced takeoff weight that can result from those benefits. All three of these characteristics would significantly reduce the amount of fuel consumed compared to current aircraft. In addition, the blended wing body has a larger internal volume that is well suited for holding alternative fuels that have a lower energy density than Jet
A 1 fuel. The larger internal volume would also allow the engines to be placed within the airframe, further reducing the drag. The design used four engines instead of two; this configuration would allow all four engines to run at takeoff and only require two during cruise.
Since engines run at maximum efficiency when run near or at full speed, the engine efficiency during cruise would be substantially increased. The inlets to the two inactive engines could then be closed off to further decrease the drag on the aircraft.
However, design issues such as a non cylindrical pressure vessel would present difficult structural problems that would need to be overcome. In addition, the blended design would necessarily increase the space needed to house the aircraft at airports, reducing the amount of room available for other aircraft. The design is also unconventional and would take considerably longer to certify compared to a more conventional aircraft design.
13
Figure 4.1 Blended Wing Body Concept
4.2 Hybrid Blended Conventional
The Hybrid Blended Conventional aircraft is an attempt to merge the industry standard cantilever wing with the Blended Wing Body aircraft design. This design is used to try to increase the overall lift, decrease the drag, and increase the fuel storage volume. The aircraft is designed with a blended fuselage wing connection that is used to decrease the amount of interference drag that is caused in this area. The blending of these two aircraft components is thought to increase the lifting surface for the aircraft to achieve greater lift. A V tail is also incorporated in this design in order to decrease the required tail surface area; with one less surface, the drag for this element is decreased [1]. The fuselage is flatter as well in an attempt to increase the internal volume to allow for more fuel to be stored. This has the implication of increasing the total amount of material required to make up the structure, increasing the weight.
14
Figure 4.2 Hybrid Blended Conventional
4.3 Strut Braced Wing
The strut braced wing design has many advantages due to its strut, which provides added support of the wingbox. Less material is required to structurally reinforce the wing due to the support by the strut, therefore decreasing TOGW by up to 10% compared to the 737. The wing weight is significantly decreased, and the decreased thickness to chord ratio permits the aircraft to cruise at high Mach numbers with a lower than conventional wing sweep angle. The decreased thickness also reduces the airplane’s overall drag by reducing the surface area of the wing. Lastly, the smaller chord, increased span, and therefore increased aspect ratio, reduce the induced and parasite drag and provide an opportunity for lower Reynolds numbers and a more laminar flow along the wing surface.
15
Figure 4.3 Strut Braced Wing Concept
4.4 Design Selection
Through the analysis and comparison of the three aircraft concepts, shown in Table 4.4.1, the final design was chosen to be the strut braced wing. The main features that contributed to the decision to use the strut braced wing design were its structural integrity, lift to drag ratio, marketability, and stability. The design receives high marks in each of these categories.
The strut braced wing receives high marks for its marketability and stability due to it having a similar style to the convention aircraft. Since this design would not be a major change to what commercial transport aircraft look like, it would be an easy sell to consumers and aircraft companies. This similarity also helps this design to score highly on the stability aspect.
16 This concept was able to triumph in the areas of structural integrity and lift to drag ratio due to its incorporation of a strut. The strut decreases the amount of force the main wing spar needs to support, reducing the shear and bending moment acting at the root chord. This in turn decreases the amount of material that is needed for this section, which improves the take off ground weight. Using the strut to carry part of the wing load, the span can be increased to obtain a lower induced drag acting on the wing. This reduction improves the lift to drag ratio for this aircraft.
Table 4.4.1 Aircraft Decision Matrix
Hybrid Blended Strut Braced Wing Blended Wing Body Weight Conventional Criteria [%] Weighted Weighted Weighted Score Score Score Score Score Score Takeoff Gross Weight (TOGW) 12 2 0.24 1 0.12 3 0.36 Internal Volume 6 1 0.06 2 0.12 2 0.12 Fuel Weight 10 2 0.2 1 0.1 3 0.3
CD0 8 1 0.08 2 0.16 3 0.24 Lift to Drag Ratio 14 2 0.28 1 0.14 3 0.42 Marketability 5 3 0.15 2 0.1 1 0.05 Cost 5 1 0.05 2 0.1 1 0.05 Manufacturing 7 3 0.21 2 0.14 1 0.07 Stability 8 3 0.24 2 0.16 1 0.08 Maintenance 5 2 0.1 2 0.1 1 0.05 Structural Integrity 10 3 0.3 2 0.2 1 0.1 System Integration 5 2 0.1 2 0.1 1 0.05 Certifiability 5 3 0.15 2 0.1 1 0.05 Totals 100 2.16 1.64 1.94
5. Sizing
5.1 Initial Weight
Analysis of a design depends heavily on knowing the weight of the aircraft. The initial design weight is essentially an estimate based on the aircraft’s characteristics and dimensions.
17 This estimate is later refined when the weights of individual components are known. Nicolai’s aircraft sizing program is used to determine these estimates.
To fully understand the use of Nicolai’s sizing program, it is used for the Boeing
737 800. Knowing the take off ground weight (TOGW) and dimensions for this aircraft, constants in the program can be changed for it to produce the appropriate weight values. This gives the constants that are used for the analysis of the design concepts. Table 5.1.1 shows the constants that are used for reserved fuel weight fraction, trapped fuel weight fraction, and the structural technology factor
Table 5.1.1 Constants from Boeing 737 800 analysis in Nicolai’s Program
Reserve Fuel Weight Fraction 0.05 Trapped Fuel Weight Fraction 0.01 Structural Technology Factor 0.84
The major factors contributing the aircraft weight for this program are the aspect ratio and the zero lift drag (C D0 ). These parameters are determined from the dimensions of the aircraft design and are determined from the following equations [15]:
(5.1.1) =
(5.1.2) =
The historical value for the skin friction coefficient Cfe is given as 0.003 for commercial
transport aircraft [1]. The values for the aspect ratio and C D0 are given in Table 5.1.2, along with the other variables for the aircraft design. The constants that are used for the weight estimation are given in Table 5.1.3.
18
Table 5.1.2 Variables for the different aircraft designs
Strut Braced Boeing 737 Wing 800
Passenger Size 175 175 Payload (lbs) 37000 47000 Mach Number 0.8 0.785 Aspect Ratio 12.23 9.45 Dynamic Pressure (psf) 223.3 214.9
CD0 0.0185 0.016 Reference Area (ft^2) 1534.44 1344.5
Table 5.1.3 Constants for Nicolai’s sizing program
SFC 0.37 Speed of Sound (ft/s) 969 Range (nm) 3500 Loiter Time (minutes) 65 Reserve Fuel Fraction 0.05 Trapped Fuel Fraction 0.01
Using the constraints listed in Tables 5.1.1, 5.1.2, and 5.1.3 above, Nicolai’s sizing program is used to determine the TOGW, weight of the fuel, and the empty aircraft weight. The
Boeing 737 800 is included in the in this analysis for comparison to a current aircraft. The weights determined from Nicolai’s program are presented in Table 5.1.4. This program does not take into account the weight savings that are gained by using a strut braced wing design. These values are a starting point for all design analysis that is conducted and will be refined with further weight estimations for individual aircraft components.
19 Table 5.1.4 Weights that are determined by using Nicolai’s Program
Strut Braced Boeing 737 800 Wing Take Off Ground Weight (lbs) 144775.54 173852.2 Fuel Weight (lbs) 48758.88 56666.7 Empty Weight (lbs) 59016.85 70185.6
5.2 Thrust to Weight and Wing Loading
Through the use of the aerodynamic data of the design and requirements from the RFP, it is possible to create a plot showing a viable thrust to weight ratio and wing loading. The RFP specifies that the landing speed must be less than 140 KCAS. FAA requirements state that the approach speed must be 30% greater than the stall speed. A landing speed of 135 knots is used to calculate the wing loading for a stall speed of 117 KCAS and a maximum landing lift coefficient of 2.4 [15]. The lift coefficient assumes the presence of double slotted flaps and slat for landing from figure 5.3 in Raymer [15]. The stall speed wing loading is calculated by assuming lift equals weight and using the stall speed and maximum lift coefficient [15]. The thrust to weight ratio as a function of wing loading for takeoff is determined by using figure 5.4 to determine the takeoff parameter (TOP) and solving for the thrust to weight ratio [15].
The density ratio is chosen to be 0.75 to account for a hot day in Denver. The TOP for a twin engine aircraft is found to be 205 lb/ft 2 [15]. The maximum takeoff lift coefficient is assumed to be 80% of the maximum landing coefficient, which equates to 1.92. The takeoff lift coefficient is assumed to be 83% of the maximum takeoff lift coefficient, which results in 1.6
[15]. To find the cruise function, equations 5.28 and 5.29 from Raymer [15] were combined and used with the aspect ratio and skin friction drag specified from the aerodynamic data.
(5.2.1)
20 The dy namic pressure for the cruise condition was calculated to be 933 lb/ft2 with the parasite drag, aspect ratio, and Oswald efficiency factor coming from the aircraft aerodynamic data. The climb gradient was chosen to be 0 since the aircraft was being analyze d in a cruise condition.
A summary of the important FAR re quirements is shown in Table 5.2.1 [15]. Figure 5.2.14 shows the thrust to weight and wing loading characteristics for the aircraft to meet FAR and RFP requirements. This allows for an optimal choi ce of wing loading and thrust to weight to be 108 lb/ft 2 and 0.33 respectively.
Table 5.2.1 FAR Requirements
Landing Minimum G Flaps Gear (%) Second Segment Takeoff Up 3 Climb (2 Engines) Go around Ldg. Landing Down 3.2 Config. (2 Engines)
Figure 5.2.1 T/W vs. W/S
21 6. Aerodynamic Performance
6.1 Drag Polar
Aircraft aerodynamic forces of lift and drag are attributed to combinations of shear and pressure forces. Total drag on the aircraft is the sum of parasite drag and induced drag.
Parasite, or zero lift drag, is estimated through a component buildup method from calculated flat plate skin friction drag and pressure drag. A component form factor is applied to estimate the influence of viscous separation on the pressure drag [15].
Induced drag, or drag due to lift, is due to circulation around a lifting body that produces trailing edge vortices. These vortices result in a force on the body that constitutes a form of drag proportional to the square of the lift coefficient. This aircraft is identified to cruise in a transonic regime with a cruise Mach number of 0.8. In transonic flight, wave drag becomes significant in the estimation of the pressure drag due to the formation of shocks [23].
Table 6.1.1 presents the total drag at takeoff, cruise of Mach 0.8 and altitude 35,000 ft, and approach for each component of the aircraft. Figure 6.1.1 displays the corresponding drag polars at cruise, takeoff, and approach conditions.
Table 6.1.1 Cruise Component Drag
Drag Coefficient Component Takeoff Cruise Approach Fuselage 0.0064 0.00632 0.0071 Wings 0.0022 0.00284 0.0022 Struts 0.00216 0.00191 0.00218 Engines 0.0017 0.00144 0.0018 Vertical Tail 0.00149 0.0012 0.00149 Horizontal 0.0016 0.00132 0.00165 Tail Flaps 0.0075 0.021 Landing 0.0198 0.0198 Gear Total 0.04285 0.01503 0.05722
22 2.5
2
1.5 L
C Cruise 1 Takeoff Approach 0.5
0 0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2
CD
Figure 6.1.1 Cruise Drag Polar
6.2 Lift-to-Drag Ratio
The lift to drag ratio (L/D) is an important aerodynamic efficiency indicator. The theoretical maximum lift to drag ratio is given by the following equation [15]:
(6.2.1) = The Oswald efficiency factor is typically between 0.7 and 0.85 for a subsonic, moderately swept aircraft [15]. The EN 1 has an Oswald efficiency factor 0.74, and using the cruise drag coefficient C D0 of 0.1503, the aircraft is projected to have an (L/D) max of approximately 21.68 for cruise. The Boeing 737 800 has a cruise (L/D) max value of
approximately 17.26. The RFP requires an improvement over the Boeing 737 by 25%, and so the
required (L/D) max value for cruise must be at least 21.58. Therefore, the EN 1 design meets this
critical design requirement. The improvement of lift to drag ratio will have a substantial effect
on the efficiency of the aircraft as a whole, and thus significantly reduce emissions and costs.
23 6.3 Airfoil Selection
The EN 1’s airfoils were selected using analysis from the modified Korn Equation:
⁄ (6.3.1) = − −
The drag divergence Mach number M dd is dependent on the technology factor , the thickness to chord ratio (t/c), the section lift coefficient C l, and the wing sweep Λ.
The primary constraints on the airfoil selection are the drag divergence Mach number and
the natural laminar flow for each airfoil. The fuel volume contained within the wing is a
secondary consideration.
The drag divergence Mach number, as calculated in the Korn Equation, must be larger
than the cruise Mach number of 0.8 so that there is no excessive drag on the wings. There is also
the desire to have a M dd that at least matches the maximum design Mach number of 0.83. This requirement is influenced by each variable in the Korn Equation. Increasing the sweep and the technology factor, and reducing the section lift coefficient and thickness to chord ratio, increases
Mdd .
The RFP states the requirement for enhanced natural laminar flow, which is best met by reducing the wing sweep and the (t/c) ratio.
The fuel contained within the wing serves to alleviate the bending moment caused by the wing creating lift, thus reducing the structural weight necessary to keep the wing intact. A larger
(t/c) ratio increases the amount of fuel that can be contained within the wings, but this would also decrease M dd .
The NASA supercritical airfoils [26] were selected as the set of airfoils to be considered.
These airfoils are designed to have high drag divergence Mach numbers, as reflected in their
24 high technology factor of 0.95 in the Korn Equation, and have a number of combinations of section lift coefficients and thickness to chord ratios to choose from.
Figure 6.3 .1 shows plots of (t/c) against values of sweep and C l for the NASA supercritical airfoils at M dd of 0.8 and 0.83. Due to the requirement of enhanced laminar flow, the sweep angle was constrained to below a value of 18 degrees for the main wing. The section lift coefficient was limited to above 0.4 due to concerns that any less would not produce enoug h lift.
Based on this plot, it is obvious that the optimal (t/c) ratio is 0.06. However, concerns about the structural integrity of the main wing with that thickness, as well as the small fuel volume, eliminated that option. For the minimum M dd , the (t/c) values of 0.10 and 0.12 meet the constraints. However, only a 10% thick airfoil is able to meet the maximum Mach number of
0.83 within the constraints. Since minimizing the (t/c) ratio is desired, the 10% thickness airfoil set was selected. From this set, airfoils with section lift coefficients of 0.4, 0.6, and 0.7 were available.
1.4
1.2 Maximum Sweep
1 t/c =0.12, Mdd = 0.8 0.8
l t/c =0.10, Mdd = 0.8 C 0.6 t/c = 0.06, Mdd = 0.8
0.4 t/c = 0.12, Mdd = 0.83 t/c = 0.10, Mdd = 0.83 0.2 t/c = 0.06, Mdd = 0.83 0
0 5 10 15 20 Minumum C l Sweep [degrees]
Figure 6.3.1 Cl vs. Sweep Angle for various values of (t/c) and M dd
25
To minimize the drag produced due to the drag divergence Mach number, three airfoil sections were selected for the main wing. The root airfoil section is the SC(2) 0710, the midspan section is the SC(2) 0610, and the wingtip section is the SC(2) 0410, with the three sections blended into each other along the span of the wing for a smooth transition between lift coefficients. The three airfoil sections are shown in Figure 6.3.2. The use of multiple airfoils sections along the span will allow the wing to maximize lift produced near the root of the wing while minimizing drag by increasing the drag divergence Mach number toward the wingtip. This will also reduce the bending moment along the wing, allowing a reduction of the wing’s structural weight. In addition, since the three airfoils are of the same family, the differences between airfoils are small, and thus the airfoils will be easier to blend together.
The sweep angle was selected to be 16 degrees. This angle is the minimum angle allowed for the SC(2) 0710 airfoil while meeting the requirement of a drag divergence Mach number greater than 0.8. The SC(2) 0410 airfoil section has a M dd of greater than 0.83 at this sweep
angle, meeting our desired criteria.
The SC(2) 0410 was also selected for the horizontal tail. This was based on the criteria
that the horizontal tail would need enough internal space for structural support; the 6% thick
airfoils were considered too thin for this requirement.
26 SC(2)-0410
0.08
t/c -0.02 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
-0.12 x/c
SC(2)-0610
0.08
t/c -0.02 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
-0.12 x/c
SC(2)-0710
0.08
t/c -0.02 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
-0.12 x/c
Figure 6.3.2 Main Wing Airfoil Sections: SC(2) 0410, SC(2) 0610, SC(2) 0710
6.4 Airfoil Analysis
The RFP states that a boundary layer transition delay of 20% for a 30 degree wing sweep should be achieved through means of natural laminar flow control (NLFC). Natural laminar flow control uses the contours of the airfoil to maintain a gently accelerating velocity distribution over the wing, thus allowing the boundary layer to remain laminar for greater percentages of the
27 chord, up to 50 60% of the chord length. The drag on an airfoil utilizing NLFC is typically two thirds of a fully turbulent airfoil. This is as opposed to active laminar flow control, which uses mechanical systems to maintain the boundary layer artificially and can achieve transition delay up to the full length of the chord in the most extreme cases, resulting in a total drag of only a ninth of a fully turbulent airfoil. Thus, laminar flow control’s greatest benefit is that of increasing the L/D of the aircraft and improving its efficiency. Natural control can only be achieved within a certain range of sweep angles and Reynolds numbers, as shown in Figure 6.4.1. [33] The EN 1 has a wing sweep of 16 degrees and a wing Reynolds number of approximately 20 million and below, putting it right at the edge of the NLFC range, but still within its boundaries. Therefore,
NLFC is possible with the EN 1.
To measure the effectiveness of the EN 1’s airfoils, a boundary layer analysis was done to determine the transition and separation locations on the NASA airfoils compared to the standard 737 airfoils. Due to the similarity in shapes between the three NASA airfoils, only the
SC(2) 0710 airfoil was analyzed for simplicity, as the results would be similar for the
SC(2) 0610 and SC(2) 0410. The Java based airfoil analysis program JavaFoil was used to perform the calculations. The program uses an integral boundary layer method to analyze the airfoils, utilizing 2nd order Runge Kutta integration with stabilization by automatic step reduction. [27] The airfoils were analyzed at cruise conditions of Mach 0.8 and a Reynolds number of 5,870,000 for a unit length airfoil. The analysis is primarily qualitative, as the analysis does not take compressible effects into account; therefore, the results are measured in terms of the airfoils relative to each other. The boundary layer analysis results are shown in Figures 6.4.2 through 6.4.4.
28
Figure 6.4.1 Limits of laminar flow control technologies [33]
As can be seen by the figures, the SC(2) 0710 airfoil performs significantly better than the Boeing 737 airfoils. The transition points for the 737 root airfoil are at 17.5% and 20% chord for the upper and lower surfaces, respectively, and the transition points for the 737 midspan airfoil are 40% and 1% for the upper and lower surface, respectively. By comparison, the
SC(2) 0710 transitions at 27.5% and 49% for the upper and lower surfaces, respectively – an average delay of 34%. This indicates a much more consistent boundary layer for the EN 1’s entire wing and shows that there is a natural laminar boundary layer over a much larger percent of the chord, on average, than the 737’s airfoils, particularly on the lower surface. This results in a lower frictional drag on the SC(2) 0710 airfoil than the 737’s airfoils, improving the lift to drag ratio. This shows that the chosen airfoils are capable of exceeding the required 20% transition delay mandated by the RFP, and the delay will be further enabled by the sweep reduction.
29 In addition to the delayed boundary layer transition from laminar to turbulent, the
SC(2) 0710 and its family of airfoils also have a much delayed boundary layer separation at cruise conditions. The Boeing 737 airfoils’ boundary layers separate from the surface after transitioning within a few percent of the chord in almost every case. In contrast, the
SC(2) 0710’s boundary layer does not separate until 81% and 99% of the chord for the upper and lower surfaces, respectively. This improvement, an average of 41%, allows lift generation to occur over a larger area of the wings than the wings of the Boeing 737, further improving the lift to drag ratio.
737 Root Upper δ1
0.008 Upper δ2 0.007 Upper δ3 0.006 Lower δ1 0.005 Lower δ2 δ 0.004 0.003 Lower δ3
0.002 Upper Transition
0.001 Upper Separation 0 Lower Transition 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c Lower Separation
Figure 6.4.2 Boeing 737 Root Airfoil Boundary Layer
30 737 Midspan Upper δ1 0.004 Upper δ2 0.0035 Upper δ3 0.003 Lower δ1 0.0025 Lower δ2 δ 0.002 Lower δ3 0.0015
0.001 Upper Transition
0.0005 Upper Separation
0 Lower Transition 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 Lower Separation x/c
Figure 6.4.3 Boeing 737 Midspan Airfoil Boundary Layer
SC(2)-0710 Upper δ1
0.01 Upper δ2
Upper δ3 0.008 Lower δ1 0.006 Lower δ2 δ 0.004 Lower δ3
Upper Transition 0.002 Upper Separation
0 Lower Transition 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 Lower Separation x/c
Figure 6.4.4 NASA SC(2) 0710 Airfoil Boundary Layer
31 7. Propulsion
7.1 Engine Technologies
The latest engine to come from Pratt and Whitney is the Pure Power PW1000G turbofan.
This engine is planned to be put into service by 2013. The PW1000G will increase engine efficiency through a clever gear system which allows the fan to operate at a slower rpm while allowing the compressor and the turbine to operate a higher more efficient rpm. With the gear box, the engine will consume less fuel, and decreases engine noise and emissions. The increase in the engine efficiency allows for fewer engine stages, which means less engine weight and a lower maintenance cost. Other components that affect the engines efficiency are aerodynamic improvements, composite materials to decrease weight, and advanced internal components. The turbofan will operate with a high pressure spool, a low pressure turbine, a combustor, advanced engine controls, and an engine health monitoring system. The health monitoring system will be used to lower maintenance cost by identifying engine defects before they become a problem. The engine specifications for the Pure Power PW1000G turbofan are in Table 7.1.1.
Table 7.1.1 Pratt and Whitney PW1000G specifications [17]
Pure Power PW1000G Thrust (lb): 14 17,000 17 23,000 Fuel Burn : 12% 12% * vs current engines Noise (dB) : 15 20 * Stage 4 CO2 Reduction : 2,700 3,000 * Tonnes per year NOx Reduction : 50% 55% * margin to CAEP 6 Fan Diameter (in) : 56 73 Weight : Less Less * vs current engines
The next generation of turbofans from GE is the 1854. This turbofan is expected to GE decrease emissions and fuel consumption by using composite materials, a clean burning TAPS
32 combustor, special internal coatings, use of counter rotating architecture, and a low maintenance fan module. Other technologies that are used by the 1854 are a 10 stage high pressure GE combustor, composite fan blades, and composite fan case. The engine weight is expected to be
less due to reducing the number of engine components by 30% and using composite materials.
The engine specifications for the General Electric turbofan are in Table 7.1.2. GE Table 7.1.2 General Electric 1854 specifications [18] GE Genx 1854 Thrust (lb): 53,200 Fuel Burn : 15% Noise (dB) : CO2 Reduction : 95% * less from regulations NOx Reduction : 95% * less from regulations Fan Diameter (in) : 111.1 Weight : Less
The next generation turbofan from Rolls Royce is the Trent 1000. This turbofan uses a 3 shaft design which lowers the engine noise by having a lower jet velocity and a slow fan speed.
A tiled combustor will be used to reduce engine emissions to a level lower than that of any large turbofan in service. The Trent 1000 incorporates soluble core high pressure turbine blades, new manufacturing methods, and new materials to improve the longevity of the internal components of the engine. The Trent 1000 uses an upgraded version of Rolls Royce’s Predictive Engine
Health Monitoring system which can now pinpoint maintenance needs before they can disrupt normal operation of the aircraft. Another feature of the Trent 1000 is the Intermediate Pressure spool power offtake which generates more electrical energy output for the aircraft than current engines. The engine specifications for the Rolls Royce Trent 1000 turbofan are in Table 7.1.3.
33 Table 7.1.3 Rolls Royce Trent 1000 specifications [19]
Trent 1000 Thrust (lb): 53 – 75,000 Fuel Burn : 15% *vs. Trent 700 Noise (dB) : 3.8 to 5 CO2 Reduction : 12 to 14 % *vs. CFM56 NOx Reduction : 40 to 50 % *vs. CFM56 Fan Diameter (in) : 112 Weight (lb) : 11,924 Length (in) : 160
The modern turbofan to come out of CFM is the CFM Leap X. This turbofan is set to debut in the year 2016. The Leap X turbofan will run on less fuel, generate less fuel, and generate less harmful emissions. The Leap X will achieve these feats with composite fan blades and a composite fan case. The engine will use 3 D woven transfer molding blades to gain more efficiency out of the turbofan as well as decrease the weight of the engine. The composite fan should take away 500 pounds off of each engine. The turbofan will produce fewer emissions NO by using a Twin Annular Pre Swirl combustor. Other features include an 8 stage compressor, a single stage high pressure turbine, and innovative cooling systems. The engine specifications for the CFM Leap X turbofan are in Table 7.1.4.
Table 7.1.4 CFM Leap X specifications [20]
CFM Leap X Thrust (lb): 18 – 50,000 * Leap X Thrust Range Fuel Burn : 16% *CFM56 Noise (dB) : 10 or 15 CO2 Reduction : 16% *CFM56 NOx Reduction : 60% *CAEP/6 regulations Fan Diameter (in) : Weight (lb) : 500 *less than CFM56 Length (in) :
34 7.2 Engine Selection
After researching multiple jet engines, the CFM Leap X engine is the best fit for the aircraft. Compared to the current CFM 56 jet engine, the Leap X has less emissions, noise, and weight, while providing adequate thrust with a reduced fuel burn. When compared to other modern jet engines, the Leap X had the lowest CO 2 and NO x emissions, does not need to be
scaled down to produce the required thrust for the EN 1, and may be certified as early as 2016.
7.3 Fuel System
The EN 1 contains three fuel tanks, two in the wings and a third constructed from aluminum alloy within the fuselage. The main fuel tank is situated below the main cabin, underneath the center of gravity. Each wing contains a fuel tank that extends from 10% of the wing span to 70% of the span. The tank occupies the space between 12% and 60% of the local chord, between the forward and aft spars, and has an allowance of approximately one inch between the wing surface and the fuel for the primary wing structure.
An Excel program was developed to calculate the fuel volume based on these parameters and is able to be added on to in a modular fashion. The program gives a unit fuel tank area for an airfoil section of 0.040354 ft 2/c 2, shown in Figure 7.3.1. The total fuel tank volume contained within the wings is 459 ft 3. Seven percent of the volume is subtracted to account for wing
structure, and an additional 3% subtracted to allow for fuel expansion as stipulated in section
25.969 of the FARs. Therefore, the total fuel volume contained within the wings is 413 ft 3. The
center of gravity for the fuel contained within each wing is shown in Figure 7.3.2, with a range
line to indicate how it will travel as fuel is burned from the wing from either direction.
35 0.2
0.15
0.1
0.05
0 t/c -0.05 0 0.2 0.4 0.6 0.8 1
-0.1 Airfoil Fuel Tank -0.15 Fuel C.G. -0.2 x/c
Figure 7.3.1 Airfoil Fuel Tank and C.G. Location
Spanwise Position [ft] 0 10 20 30 40 50 60 70 0
5 Wing Outline Fuel Tank 10 Fuel C.G. Line 15
ChordwisePosition [ft] 20
25
Figure 7.3.2 Wing Fuel Tank and C.G. Location
7.4 Engine Maintenance
An important aspect to consider when designing an aircraft is the accessibility of the engine for maintenance. Easy access ensures decreased maintenance time, saving the airline company money. For the EN 1, each side of the nacelle is designed to open upward on hinges, as seen in Figure 7.4.1 to allow for maintenance while the engine is still on the aircraft. The engine
36 can be removed from the plane completely for an engine swap from the front of the nacelle. With the side panels up, a dolly will pull the engine out and a new one will replace it. The engine sits far enough forward that its maintenance and removal can be conducted without interference with the wing or strut.
Figure 7.4.1 Nacelle panels open for maintenance and engine removal
8. Performance
8.1 Takeoff, Landing, Balanced Field Length Analysis
A summary of the takeoff, landing, and balanced field length results are displayed in
Table 8.1.1. The takeoff analysis is broken into 4 different segments for separate analysis followed by a summation of all the ground distances to compute the total takeoff distance. First, the ground roll is calculated as the distance from zero velocity until the point of rotation, which occurs at V TO . The takeoff velocity is assumed to be 1.1V stall . The average thrust over this distance is assumed to be the thrust at 70% of V TO . The ground roll is calculated as an integral
over the varying ground roll velocity of the velocity and acceleration. The second segment is the
transition distance which is assumed to take 3 seconds for an airliner. Since the acceleration is
negligible in this portion, the ground distance traveled is three times the takeoff velocity. The
third segment is the transition to climb segment with an assumed average velocity of 1.15V stall .
37 The average lift coefficient during transition is assumed to be 90% of the maximum takeoff lift coefficient of 1.92. A vertical load factor is used along with the transition velocity to calculate a radius of curvature to a best climb angle of 12.9 degrees. The ground distance is then calculated using trigonometry. The height gained during transition is also calculated to determine if the
50 ft. obstacle has been cleared as specified in the FAA Part 25 regulations. In this case, the airline cleared the obstacle during transition.
Table 8.1.1 Takeoff, Landing, and Balanced Field Length Results
Takeoff Distance 5678 ft Balanced Field Length 9093 ft Landing Distance 7370 ft
The landing analysis is performed in a similar manner, but in reverse of the takeoff analysis. Using a 50 ft. obstacle to clear, an approach speed of 150 knots, and a 3 degree approach, an approach distance was calculated. A flare distance was calculated like the transition for takeoff using a flare speed of 1.23V stall and a load factor of 1.2. A brake delay after touchdown is assumed to be 2 seconds traveling at a velocity of 1.15V stall . This results in a free roll distance calculated as the brake delay multiplied by the touchdown velocity. The ground roll is then calculated the same as the takeoff with the thrust being the idle thrust and the initial velocity being the touchdown velocity. No thrust reversers were used in the calculation since the
FAA does not allow them to determine landing distance in case of malfunction.
The balanced field length is calculated from an equation summarized in Raymer [15]. It involves the use of the climb gradient required by the FAA, wing loading, obstacle height, average thrust, air density, and climbing lift coefficient. The RFP specifies that the takeoff distance be less than 8,200 feet. This aircraft far exceeds the specified requirements.
38 8.2 Mission Profile
The mission segments and corresponding range and fuel burn is shown in Table 8.2.1, while Figure 8.2.1 shows the mission profile segments. The mission profile is split into 11 segments with 7 of the segments involved in the primary mission of the aircraft. The final 4 segments are used to meet the FAR requirement for 45 minutes of additional fuel and the capability to divert to another airport.
The fuel burn for each segment was calculated by using methods described in Raymer
[15]. Specifically, estimates for Warm up/Takeoff, climbing, and landing were adopted from
Raymer [15]. The fuel burn during cruise was approximated using the breguet range equation.
The loiter segment was calculated by using an endurance equation with a loiter time of 45 minutes while flying at the maximum lift to drag ratio [15]. To be conservative, the descent segments 4, 6, and 10 used a fuel fraction of 0.99, 0.995, and 0.999 respectively. The fuel fraction is the ratio between the initial weight during that segment and the final weight at the end of the same segment. The mission profile analysis shows a fuel weight of about 40,000 lbs to fly a maximum range of 3,900 nm. Including the diversion range, the aircraft can fly about 4,200 nm.
Figure 8.2.1 Mission Profile
39 Table 8.2.1 Mission Segment Analysis
Range Fuel Burn Mission Segment (nm) (lbs) 1. Warmup/Takeoff 0 4338 2. Climb 35 2104 38 3. Cruise 00 25063 4. Initial Descent 50 1131 5. Loiter 0 2022 6. Final Descent 20 550 7. Land 0 547 8. Climb 20 1633 25 9. Cruise 0 1403 10. Descend 20 106 11. Land 0 529 41 Totals 95 39819
9. Weights and Structures
9.1 Final Weight
One of the integral parts of the conceptual aircraft design process is weight estimation, and its importance is evident even as the design progresses. Initial sizing methods provide crude estimates of takeoff weight and fuel weight required for particular flight missions, and as previously stated they are a starting point for all design analysis.
The following method involves estimating weights of individual components, and their sum is an estimate of the aircraft’s total weight. Utilizing this method allows one to easily calculate a rough estimate of a conceptual aircraft’s center of gravity location. The c.g. of an aircraft is the point at which its weight is concentrated and where the aircraft would balance. It
40 is also the point about which the aircraft rotates due to forces generated by its control surfaces.
Engineering groups rely on a c.g. estimate early in design to avoid additional work later in the design process after a more accurate c.g. is properly estimated. This method’s crude component buildup technique involves planform and wetted areas and gross weight percentages, which is suitable for preliminary balance calculations. Using Table 15.2 in Raymer [15], one can calculate various aircraft component weights. Based on the statistical approach and historical values for the weight per unit area of exposed planform area, the weights of the wing and tail are calculable. Similarly, the fuselage weight is determined using its wetted area. The installed engine weight is a multiple of the uninstalled engine weight, and other components, such as landing gear, is estimated as a fraction of the takeoff gross weight [15]. These estimated weights were used to verify more detailed calculations and statistical equations that were used to solve for components’ weights of the EN 1, Equations 15.25 through 15.44 in Raymer [15].
The more detailed, statistical equations used for weight calculations of EN 1 components are based upon existing aircraft data. In order to account for non conventional designs such as braced wings and advanced composite structures, adjustment factors determined by Raymer were used. The adjustment factors are multiplied by the component weights to estimate the novel designs’ components’ weights. The final weight based on the detailed, statistical method is utilized in the final aerodynamic, stability and control, and cost analysis. The initial sizing outputs for TOGW and empty weight for the strut braced wing, 144,800 lbs and 59,000 lbs, respectively, are comparable to the results of the detailed, statistical equations. The final takeoff gross weight of the strut braced wing aircraft with advanced composites calculated using the detailed, statistical equations is 144,600 lbs with an empty weight of 49,800 lbs. The weight savings in the latter component weight calculation method come from the use of advanced
41 materials throughout the aircraft. The significant weight savings in comparison to the Boeing
737 800 come from the combination of implementing a strut braced wing and advanced materials. Figure 9.1.1 compares the different components’ weights for both advanced composites and conventional aircraft materials. Conventional materials and no adjustment factors are considered for components that depict no weight savings because there presently have been no projected improvements for such components.
Component Weight Comparison
20 Advanced Composites 19 Conventional Material 18 Y-Axis Ref Number : Component 17 1 : Wing 16 2 : Horizontal Tail 15 3 : Vertical Tail 4 : Fuselage 14 5 : Main Landing Gear 13 6 : Nose Landing Gear 12 7 : Nacelle Group 8 : Engine Controls 11 9 : Pneumatic Starter 10 10 : Fuel System 9 11 : Flight Controls 12 : APU Installed 8 13 : Instruments 7 14 : Hydraulics 6 15 : Electrical 5 16 : Avionics 17 : Furnishing 4 18 : Air Conditioning 3 19 : Anti-Ice 2 20 : Handling Gear 1
0 0.5 1 1.5 2 Weight (lb) 4 x 10
Figure 9.1.1 Component Weight Comparison
42 Table 9.1.1 Weight Statement
Weight, Loc., Moment, Weight, Loc., Moment, Component lb ft lb ft Component lb ft lb ft Wing 11818 63 744553 Hydraulics 204 59 12036 Horizontal Tail 880 130 114400 Electrical 577 56 32295 Vertical Tail 850 122 103700 Avionics 1235 4 4941 Fuselage 15088 60 905250 Furnishings 1924 63 121212 Main Gear 3914 61.1 238950 Air Conditioning 2579 45 116055 Nose Gear 781 19 14839 Anti Ice 294 55 16170 Nacelle 1626 55.5 90243 Handling Gear 44 5 220 Total Weight Engine 6400 55.5 355200 Empty 49756 60.4 3004743 Engine Control 54 55 2970 Fuel 55000 59.7 3283500 Starter 147 53 7799 Operating Weight 104756 60 6288243 Fuel System 303 60 18190 Crew/Passengers 31860 61.5 1959390 Flight Controls 74 100 7400 Cargo 8000 63 504000 Takeoff Gross APU install 748 130 97240 Weight 144616 60.5 8751633 Instruments 216 5 1080
9.2 Center of Gravity
After weights are estimated, one can estimate the c.g. locations of individual components, which are listed above in Table 9.1.1. The sum of components’ distances from a reference point to their c.g. locations multiplied by their respective weights equals the distance of the complete aircraft’s c.g. location times its total weight. In other words, the c.g. location is the sum of the moments of individual components about the forward tip of the aircraft divided by the total weight. Knowing that the total weight is the sum of the aircraft’s components, the c.g. location of the aircraft can be determined. In addition to verifying more detailed calculations throughout the design process, the c.g. is important for stability and control analysis. In addition to components’ c.g. locations, Table 9.1.1 lists the c.g.’s for the aircraft’s empty weight condition, operating weight load conditions, and takeoff gross weight. Figure 9.2.1, a weight c.g. excursion diagram,
43 depicts the c.g. range in percent mean aerodynamic chord while remaining within the forward and aft c.g. limits. The plot requires knowledge of fuel location and burn off data and landing gear position. The four conditions, empty weight, aircraft with fuel, aircraft with fuel and its payload, and aircraft without fuel, but with its payload, satisfy the stability requirement of a positive static margin, or a c.g. forward of the neutral point, by at least 5%, which is discussed later in the report. Lastly, the EN 1 has a c.g. range of about 7.8% mean aerodynamic chord.
140000 TOGW (129,600 lbs)
120000 Payload + Fuel
100000 Fuel Payload - Fuel 80000
60000 Weight Weight (lbs) Payload 40000
20000 Aft c.g. Limit 0 -10 0 10 20 30 40 50 60 % MAC
Figure 9.2.1 Weight C.G. Excursion Diagram – Clockwise Path
9.3 Materials
The significant weight savings of the EN 1 in comparison to the Boeing 737 800 come from the combination of implementing a strut braced wing and utilizing advanced materials. The strut braced design relieves loads in the wing, allowing for reduced wing thickness and therefore reduced structural weight. Advanced materials with high strength to weight ratios are appropriate to use for various components, but cost is an important consideration. Figures 9.3.1, 9.3.2, and
44 9.3.3 below are graphical representations of various advanced aircraft material properties. The aircraft materials compared in the figures include aluminum alloy, carbon fiber reinforced polymer laminate, glass fiber reinforced polymer laminate, steel, and titanium, and they are compared to each other based on cost per pound, relative density, and relative yield stress [28].
60 50 40 30
Cost, $/lb Cost, 20 10 0 Aluminum CFRP GFRP Steel Titanium Alloy Laminate Laminate Material
Figure 9.3.1 Aircraft Materials’ Cost per Pound
6 5 4 3 2
Relative Relative Density 1 0 Aluminum CFRP GFRP Steel Titanium Alloy Laminate Laminate Material
Figure 9.3.2 Aircraft Materials’ Relative Density
45 12 10 8 6 4 2 Relative Relative Yield Stress 0 Aluminum CFRP GFRP Steel Titanium Alloy Laminate Laminate Material
Figure 9.3.3 Aircraft Materials’ Relative Yield Stress
Figure 9.3.4 depicts the EN 1 and where particular materials are used in the aircraft’s construction. Much of the aircraft is built from composites which do not corrode or fatigue like metals. The use and locations of advanced materials on the EN 1 is largely based on its implementation on the Boeing 787 Dreamliner, which take into consideration material properties such as their susceptibility to various loads and cost [29]. In particular, engine pylons are constructed of aluminum, steel, and titanium alloys which can carry large loads from the engines, they resist fatigue, and they maintain structural integrity at high temperatures. Also, control surfaces and nacelles are constructed of a carbon sandwich composite which consists of two face sheets of either aluminum or fiberglass with graphite epoxy that carry tension and are bonded by a honeycomb or foam that carry shear loads and compression [15]. Again, Figure ___ shows where different materials are used to construct the EN 1, which saves significant weight and requires less maintenance, which saves fuel and money.
46
Figure 9.3.4 Materials Used In EN 1 Body
9.4 V-n Diagram
Structural analysis of the aircraft structure can only take place once the load factor is determined. This is accomplished by using a V n diagram for both gust and maneuver. The maneuver plot is constructed using the limit load factors for a transport aircraft of 2.5 and 1. The gust plot is created using gusts of 25, 50, and 66 feet per second. Using the performance characteristics and aircraft dimensions, the V n diagram shown in Figure 9.4.1 is generated.
47 3
2.5 Maneuver 2 Gust 1.5
1
0.5 LoadFactor, n 0 0 100 200 300 400 500 600 -0.5
-1
-1.5 Velocity, Knots
Figure 9.4.1 V n Diagram for maneuver and gust
This figure shows that the greatest load factor occurs at a gust of 50 feet per second. The load factor at this point is 2.6. It lies outside of the maneuver box due to the wind effects working on a longer wing span. This factor is used in combination with a factor of safety of 1.5, which gives a total load factor of 3.9 for all structural analysis.
9.5 Structural Analysis
The EN 1 is designed to take advantage of new composites for a large number of its structural components. Composites such as carbon fiber reinforced plastic (CFRP) are currently being used in the aviation industry to reduce the structural weight of an aircraft while still providing structural integrity. The EN 1 also utilizes a strut to support the wing. This reduces the bending moment acting on the wing root which allows for less material to be used for the wing spar and skin.
48 The wing box for the cantilever and strut braced wing is designed to have two spars with ribs spaced evenly across the span. The front spar starts at 12% chord and the rear spar is at 60% chord. The height of both spars is 7.78% chord. For simplification, the wing box is assumed to have a rectangular cross section, with the top and bottom of the airfoil considered flat.
The strut being used also has a two spar design. Both spars have a height 7.39 inches. The front and rear strut spars attach to the front and rear wing spars respectively by a jury strut and vertical offset. The wing and strut configuration is shown in Figure 9.5.1.
Figure 9.5.1 Layout of the wing, strut, jury strut, and vertical offset
A jury strut connects the strut to the wing at 33% span. It is used to prevent buckling of
the strut under compression when the aircraft experiences a negative load factor. It is also has
two spars and is discontinued in the middle to allow the engine nacelle to be placed under the
wing. Structural supports around the engine are used to connect the top and bottom of the jury
strut. A vertical offset is used to connect the end of the strut to the wing at 81.3% span. It has a
two spar design with length of 3.33 feet. It forms a 90 degree angle with the wing and a 99.66
degree angle with the strut. This offset reduces the interference drag that would be caused by
49 connecting two aircraft members. The overall wing and strut configuration can be seen in
Figure 9.5.2.
Figure 9.5.2 Overall layout of the strut and wing design
The design being used to accomplish the RFP requirements is a high wing supported by a strut. This design allows for a reduction in material used for the wing spar and allows for the wing to have a larger span. In order to see the benefits of this wing design, it needs to be compared to a cantilever wing. For the comparison, the dimensions and shape for the wing remains the same and two distributed load cases are used. The first includes only the lift and wing weight distributed loads while the second includes these as well as the fuel weight distributed load. The first case, shown in Figure 9.5.3 is more critical because without the fuel weight pulling down on the wing, the overall upward load will be greater. The second case, shown in Figure 9.5.4, is conducted to ensure that the wing’s structure will hold up in flight with the wing fuel tanks full.
50
Figure 9.5.3 Distributed load with empty wing fuel tanks
Figure 9.5.4 Distributed load with full wing fuel tanks
The wing loading due to lift is based on an equation created using the MatLab Tornado program while the wing weight and fuel loading are based on analytical modeling [24]. To
simplify the curves in Figures 9.5.3 and 9.5.4, the distributed load is approximated as being
51 linear between every two nodes. The nodes are chosen at the wing root, jury strut, strut off set, and the wing tip. These nodes are represented as black circles in the two graphs and the linear distribution is represented as red lines. All loads take into account the 3.9 load factor.
The cantilever wing is examined using traditional structural analysis. The shear force acting on the wing is determined by differentiating the distributed loads and including the force from the weight of the engine. The bending moment is calculated by differentiating the shear force and including the bending moment caused by the engine. Knowing that the maximum bending moment occurs at the wing root, this moment is used in the shear flow equations. These equations are used to determine the wing spar and skin thicknesses. Using these thicknesses, the minimum rib spacing to prevent skin buckling is computed. A MatLab code is utilized to conduct all of the necessary calculations and output the appropriate values.
Under the no fuel condition the root bending moment is 5,317,335 pound force feet. With this bending moment the required thickness of the wing spars and skin are 4.33 and 1.27 inches respectively. With this skin thickness the wing needs to have a rib spacing of 39 inches to prevent buckling. Using these dimensions with the case that includes the fuel weight, the margins of safety improve. Thus these wing box parameters are acceptable for all cases.
With the cantilever wing analysis complete for a comparison, the strut braced wing analysis is conducted. This is accomplished by using matrix structural analysis. For this method
7 nodes are created on the wing and strut structure as seen in Figure 9.5.5. The wing structures for elements 1 2, 2 3, 2 6 and 3 7 are frames, element 3 4 is a beam, and elements 5 6 and 6 7 are trusses. These elements and their structural types are also shown in Figure 9.5.5. At each node, the degrees of freedom are determined and are shown in Figure 9.5.6. For simplification purposes, nodes 6 and 7 are assumed to be pin joints, thus allowing rotation at these points.
52
Figure 9.5.5 Node placement and structural components
Figure 9.5.6 Degrees of freedom for the strut braced wing
With the nodes and degrees of freedom determined, the restrained stiffness matrix is created for the entire structure. The fixed end action matrix is created and used with the stiffness matrix to compute the nodal displacements. The stress matrix is created and used with these displacements to calculate the shear stress and bending moment. The bending moment at the wing root is utilized to determine the wing spar and skin thicknesses. These thicknesses are used to calculate the minimum rib spacing to prevent buckling.
53 From the matrix structural analysis of the no fuel scenario, the wing root bending moment is determined to be 2,556,615 pounds force feet. This is significantly less than the bending moment that would act on a cantilever wing. With this moment reduction, the spar and skin thickness can be reduced to 1.0 and 0.80 inches respectively. The thickness of the spars and skin of the strut and vertical offset need to be 0.50 inches each. The jury strut which needs to support the engine needs to be 1.0 inches. Using these dimensions for the wing structure, the ribs need to be spaced every 24.0 inches to prevent buckling to the skin. These wing dimensions are also used in the analysis of the scenario with full wing fuel tanks. For this second case, the margins of safety for the wing box structure increase. Since the design can withstand the moments from the extreme situations, it will be appropriate for all loading cases that the aircraft will experience.
The overall structural layout of the wings and fuselage are seen in Figure 9.5.7. This figure shows the placement of the spars and ribs in the wings and horizontal tail. The ribs are spaced to the 24 inches that is determined from the buckling evaluation. There are eight bulkheads used in the structural design. Pressure bulkheads are use at the front and rear of the cabin, while others are used at other key locations. One is used on each side of each exit door and there are two used at the wing attachment point. One bulkhead attaches to each the front and rear wing spars. They ensure the structural integrity of the aircraft from the wing and landing gear forces. Longerons, stringers, and frames are also used throughout the structure to carry the forces acting of the fuselage. The frames are spaced 20 inches apart as determined as the standard for a Boeing 737 replacement [32].
54
Figure 9.5.7 Structural layout for the EN 1
The bending moment and wing box dimension comparison is represented in Table 9.5.1.
From this table, it can be seen that the root bending moment is decreased 51.92% and 56.33% for the empty fuel and full fuel tank conditions respectively by using the strut braced wing. This is a drastic decrease in the bending moment acting on the root. With this moment decrease, the spar and skin thickness is decreased 76.91% and 37.01% respectively. These values support the fact that the take off gross weight is decreased by using a strut supported wing.
55 Table 9.5.1 Cantilever and strut braced wing comparison
Cantilever Cantilever SBW SBW Percent Wing with Wing with with No Fuel with Fuel Decrease No Fuel Fuel Wing Structure Weight 12251 10045 12251 10045 18.01 5317335 2556614 51.92 Root Bending Moment 5231280 2284699 56.33 Spar Thickness 4.33 1.00 4.33 1.00 76.91 Skin Thickness 1.27 0.80 1.27 0.80 37.01 Strut Spar Thickness 0.50 0.50 Strut Skin Thickness 0.50 0.50 Jury Strut Spar Thickness 1.00 1.00 Jury Strut Skin Thickness 0.50 0.50 Rib Spacing 39 24 40 26
10. Stability and Control
10.1 Longitudinal Stability Analysis
The EN 1’s stability and control aspects were analyzed using JKayVLM, a standalone executable program, [34] and Tornado, a MATLAB based program. [16] Both are Vortex Lattice
Method (VLM) codes. The programs are intended for linear aerodynamic conceptual wing design. The programs do not directly account for compressible flow that would be encountered at cruise speeds for a commercial transport aircraft due to using VLM for its analysis. However, the
Prandtl Glauert compressibility correction is implemented in Tornado to provide a simple correction for such effects. Both programs are capable of calculating aerodynamic and control surface derivatives, and Tornado can produce pressure coefficient, spanload and shear and bending moment distributions, as well as other information. [16]
56 10.1.1 JKayVLM Analysis
Using the program JkayVLM.exe, the aircraft was analyzed for longitudinal stability characteristics. Table 10.1.1 is a list of the input parameters used in the program.
Table 10.1.1 Numeric input parameters for the JkayVLM.exe program
Mach Number 0.8 Wing Area 1534 ft 2 Wing Mean Chord 11.93 ft Wing Span 137 ft Height Above Ground 11.64 ft X cg 65.0 ft Zcg 0.0
Using the above data and using an appropriately modeled aircraft, the stability characteristics results are shown in Table 10.1.2.
From the results in Table 10.1.2, a static margin of 5.2% was produced. Since the static margin is positive, the aircraft is considered statically stable in pitch. The results agree with C mq and C Lq. C mq should be negative and C Lq should be positive for an aft tail configuration. The C lβ from the results is negative, which means that the aircraft is statically stable in roll. [15]
Table 10.1.2 Output from JkayVLM.exe
CLα 9.57 Cmα 59.10 CLq 109.6 Cmq 941.1 Cyβ 0.1456 Cnβ 0.0402 Clβ 0.0089 Cyr 0.1031 Cnr 0.0308 Clr 0.00575 Clp 0.833 Cnp 0.5551
57 10.1.2 Tornado Analysis
The EN 1’s neutral point was determined in Tornado through an iterative process of specifying a reference point and calculating C mα at that point, as Tornado does not have a
dedicated process for automatically finding the location at which C mα is approximately equal to
3 zero. The margin of error for the value of C mα was set at 10 ; further refinement of the value would not change the neutral point location more than an inch in the analysis, which would be an insignificant change in the conceptual design phase.
The neutral point location was determined to be at 20.4% of the mean aerodynamic chord, 61.63 feet from the nose, shown in Figure 10.1.1. The static margin was therefore determined to be 13.4% at the furthest forward c.g. location and 5.1% for the furthest aft c.g. location. A typical transport aircraft has a static margin between 5 10% when the c.g. is furthest aft, and so the aircraft has an acceptable amount of static stability in pitch. [15]
The analysis output a set of stability derivatives for the aircraft, shown in Figure 10.1.2.
As noted above, C mq should be negative and C Lq should be positive for an aft tail configuration, and C lβ should be negative for static stability in roll. Each is shown to be true for the aircraft
configuration used in the Tornado model. This analysis shows agreement between JKayVLM
and Tornado, ensuring that the aircraft does have the necessary amount of static stability.
58 20
15
10
5
0
-5
-10 Aircraft body y-coordinate
-15
-20 0 5 10 15 20 25 30 35 40 45 Aircraft body x-coordinate
Figure 10.1.1 Aircraft Layout – C.G. is black, Neutral Point is red, MAC is blue
Table 10.1.3 Stability Derivatives from Tornado Analysis
CL CD CY
CL α 6.79E+00 CD α 1.35E 01 CY α 3.85E 07
CL β 2.30E 05 CD β 5.24E 06 CY β 2.87E 01
CL P 2.36E 07 CD P 2.80E 06 CY P 2.32E 01
CL Q 2.11E+01 CD Q 1.18E 01 CY Q 2.17E 06
CL R 1.66E 07 CD R 8.66E 09 CY R 2.81E 01
Roll Pitch Yaw
Cl α 1.10E 07 Cm α 3.05E+00 Cn α 2.95E 07
Cl β 7.52E 02 Cm β 3.61E 05 Cn β 1.44E 01
Cl P 7.71E 01 Cm P 1.48E 06 Cn P 1.43E 02
Cl Q 1.66E 06 Cm Q 8.59E+01 Cn Q 2.77E 06
Cl R 5.64E 02 Cm R 1.01E 06 Cn R 1.63E 01
59 10.2 Control Types
The aircraft will be using the three conventional control surfaces for transport aircraft.
These include the ailerons for roll, elevators for pitch, and rudder for yaw control. The ailerons are placed on the trailing edge of main wing from 50 to 90 percent span. The aileron chord over wing chord ratio is 0.24. The elevators will be placed on the tail of the aircraft and will run from
20 to 90 percent chord. The elevator chord over wing chord ratio is 0.25. The rudder will be placed on the trailing edge of the vertical tail from 20 to 90 percent chord. The rudder chord over vertical tail chord ratio will be 0.50 in order to avoid rudder effectiveness issues. Control sizing for the ailerons, elevators, and rudder were approximated using historical guidelines from
Raymer. [15]
Ailerons
Figure 10.2.1 Main wing ailerons, flaps and leading edge slats
10.3 Cruise Trim
The EN 1 was trimmed at its initial cruise conditions: Mach 0.8, altitude of 35,000 feet, and approximately 5% of its initial TOGW lost due to fuel burn. The Tornado program does not have a reliable automatic trimming procedure, so an iterative process was used to change the aircraft angle of attack and the horizontal tail incidence angle concurrently to arrive at a solution that reduces the pitching moment to approximately zero and balances the lift produced with the weight. The horizontal tail is designed to have a mechanism to alter its incidence angle in flight
60 for trim corrections. The horizontal tail incidence angle was chosen for stability correction as opposed to the elevator deflection angle because elevator deflections incur a larger drag penalty.
The primary constraint in this process was the angle of attack: changing the angle of attack alters the total lift more than changing the horizontal tail incidence angle. In addition, the angle of attack cannot be a large negative or positive value due to constraints on what it tolerable for the passengers and the flight attendants through the cruise portion of the flight. A maximum value of 4 degrees was set for the angle of attack.
The aircraft was trimmed to an angle of attack of 1.25 degrees and a horizontal tail incidence angle of 1.35 degrees. These angles have associated lift and pitching moment values that are accurate to an acceptable margin of error expected from the simplified VLM analysis.
11. Systems
11.1 Cabin Layout
The aircraft cabin layout is an important aspect of the airplane design, and is shown in
Figure 11.1.1. Based on the RFP requirements, the 175 passengers are classified into one class only. Hence, there are no divisions among the economical, business and first class passengers.
The seat with and leg space are also kept uniform. There are 30 rows with 6 seats in each row
(With the exception of the last and the first row). There are 6 lavatories with two lavatories placed in each row for a total of 3 rows. The back of the cabin is reserved for a kitchen and in flight food storage. Six attendants are desired to make the in flight experience run smooth.
Hence, these six attendants will be divided into three attendants per aisles for the total of two aisles. There are two separate restricted stairways that allow access to the luggage and storage area placed beneath the passenger seats. These mini stairways are located near the middle row of
61 lavatories. The seats for the attendants are spread around the cabin in such a way that each attendant will be seated near the emergency exits. The cabin also allows four emergency exits placed in the front and the back of the cabin.
Figure 11.1.1 Cabin Layout
11.2 In-flight Systems
The in flight systems of the strut braced aircraft model will include all the latest technologies that are available in the aircraft industry. Such systems and components are compatible with Boeing 737 and Airbus A320, which allows the systems to be compatible with the proposed strut braced aircraft model without a hitch. In flight entertainment systems from
Panasonic Avionics corporation allows passengers to access broadband and live TV with ease, while the high speed data satellite communication system (SAT 2100) allows an efficient GPS accessibility. With the help of the high resolution video cameras placed at the vital points around the aircraft, the pilots will be able to perform aircraft ground maneuvers with ease.
Many of the basic system components have been taken from the existing aircrafts in the same league due to their technologies being the latest available efficient systems in the market.
62 Such systems involve the Air Conditioning/Pressure/Ventilation, Lighting, Oxygen, Electrical wires and harnesses, fire/ice/rain protection, indicators/recorders and more. A complete list of such systems is given below in Table 11.2.1.
Table 11.2.1a List of the systems on board for the Strut Braced Aircraft
S.No. Systems Type Weight Power Supplier Rockwell 1 Communication HFS 900D high speed data radio 27.7 lbs 115 V ac, 3 phase Collins 27.5 V dc, 7.5A Rockwell VHF 920 digital data radio 11.0 lbs max max transmit, Collins SAT 2100 high speed data 13.2 kg (29 Rockwell 28 VDC satellite communication system lbs) Collins 115 VAC (96–134 10.5 lbs (4.76 IntuVue advanced weather radar VAC) 360 Hz– Honeywell kg) 800 Hz ADF 900 (Automatic Direction 3.402 kg (7.5 115 VAC, 400 Hz Rockwell 2 Navigation finder) lb) 0.3A Collins DME 900 distance measuring 5.94 kg (13.1 Rockwell 115 V ac, 400 Hz equipment lbs) Collins 115 VAC, 400Hz, Rockwell GLU 920 multi mode receiver 4.3 kg (9.5 lbs) 42 Watts Collins Hamilton 3 APU APS 3200 Auxiliary Power Unit. 308 lbs Sundstrand Max. Load: 4 Landing Gear Dunlop Aircraft Tires Dunlop 46700 lbs Crane Landing Gear Control
Interface Unit Goodrich Carbon Brakes Messier Tire pressure monitoring system Bugatti Messier Nosewheel steering system Bugatti Messier Landing gear Dowty Crew & passenger oxygen Avox 5 Oxygen system components Systems Interior Lighting System (PSU, Goodrich 6 Lights Emergency lighting, warning Interiors lights) ECE Zodiac External Lighting Systems Systems 7 Doors Passenger and Emergency Doors Latecoere
63 Table 11.2.1b List of the systems on board for the Strut Braced Aircraft
S.No. Systems Type Wt. Pwr. Supplier Air Liebherr Conditioning/ Air Conditioning Equipment: Air management system, 8 Aerospace Pressure/ including bleed air & air conditioning Lindenberg Ventilation Avionics ventilation systems Technofan Nord Micro Cabin Pressure Control Systems: Automatic cabin AG & Co. pressure control system OHG Nord Micro Ventilation control systems AG & Co. OHG Goodrich Sensors/Transducers: Pitot probe; ice detectors; air data Sensors & 9 Flight Controls total air temperature sensors Integrated Systems Liebherr Aerospace Wing Flaps: Primary & secondary flight controls Lindenberg GmbH Diehl Engine Controls: Versatile electronic control box Aerospace 10 Electrical Wire Harnesses: Wiring harnesses Labinal HiRel Electrical & Electronic Connectors: Connectors Connectos Static Dischargers HR Smith Diehl Airborne Electrical Power Supplies: Emergency power Luftfahrt
supply; static inverters Elektronik GmbH Leoni Cable Assemblies: Cable harnesses for antenna, landing Special
gear, vertical tail plane, cabin, galley Cables GmbH Pacific Power Conversion Equipment: AC power converter Scientific Artus Ice & Rain Ice Protection/Prevention Equipment: Bleed air anti ice 11 Dukes Inc. Protection systems Pacific Fire Fighting/Detection Systems: Fire suppression 12 Fire Protection Scientific systems Aerospace
64 Table 11.2.1c List of the systems on board for the Strut Braced Aircraft
S.No. Systems Type Wt. Pwr. Supplier Goodrich Indication and Sensors & 13 Fuel Quantity Indicators: Fuel measurement system Recording Integrated Systems Goodrich Sensors & Angle of Attack Indicators: Angle of attack sensor Integrated Systems Goodrich Sensors & Stall Warning System Integrated Systems Pressure Indicators: Pressure gauges QED Inc. Diehl Electronic instrument system Aerospace Acorn Maintenance 14 Metal Finishing/Polishing: Finishing surface System Technology Automated Test Equipment: Automated test equipment; EADS Test
ground test system & Services Water and CEF 15 Water Systems: Portable water system compressor Waste Industries Water Heaters: Innerline heaters for waste and water Cox and
system freeze protection Company Water Purification Systems: Versa Pure AC 2 drinking General
water microfiltration system Ecology MT Lightweight potable & waste water tanks Aerospace AG
Here, the table shows the various components for each systems division. The components are also accompanied by their weights, power requirements and the supplier’s name. The placement of these systems has been done in the same way as the existing aircrafts in its segment. Since the cabin size and layout is very similar to the existing aircrafts, the most efficient and the reasonable locations has been sorted out already in the existing Boeing 737’s or the Airbus A320’s.
Keeping safety in mind while explaining the systems of this aircraft design, it should be noted that the aircraft cockpit is equipped with the highest safety and security instruments. These include surveillance cameras placed outside the cockpit door and are spread around the cabin for
65 the pilots to monitor the safety and respond immediately to any possible threat encounters. The communication systems use new methods for encrypting the conversations between the pilots and the traffic control so that they do not get leaked.
11.3 Cockpit Systems
The cockpit firewall section, displayed in Figure 11.3.1, includes LCD displays for aircraft controls, flight navigation, and weather information. Both captain and first officer have equal amounts of displays for aircraft control. Backup instruments have been included for safety requirements of flight. Standard controls such as the yoke, throttle, spoiler lever, rudder, and braking system are present for both seats of the cockpit. The flight navigation computers are used for upload and download of air traffic and airport information and flight planning.
3
3 6 7 6 3 1 5 5 1 1 2 1 4 2 5 5
10 1 8 1 10
9 9
1 – Multi function display areas with soft menu controls 2 – Side displays 3 – Auto pilot controls 4 – Center display 5 – Backup instrument indicators 6 – Flap and gear control levers 7 – Instrument display 8 – Throttle and spoiler control levers 9 – Flight navigation computer 10 – Yoke, rudder, and brake controls
Figure 11.3.1 Cockpit Systems Layout
66 11.4 Ground systems
The structure of the strut braced aircraft is nearly similar to the existing aircrafts such as
Boeing 737 or the Airbus A320. Hence, the ground systems such as the baggage deployment machines, passengers’ skywalk, fuel injecting systems and food loading systems will remain the same. This is a very import positive point for this design as no capital will be required to replace or build new ground systems for this particular aircraft design.
The strut braced design has a larger wingspan than the regular Boeing 737. However, it is not larger than the bigger aircraft categories such as the Boeing 747 or airbus A360. Hence, this design can be accommodated, maneuvered into hangars and airports and is compatible with the ground systems of the airport which can accommodate Boeing 747’s or larger aircrafts.
However, its larger wingspan can hinder its compatibility with the existing Boeing 737 ground controls and hangars (if the airport cannot accommodate aircrafts larger than the size of Boeing
737’s).
Fuel transportation and storage is a major concern here. Since, the desired fuel for this design is SPK, new fuel processing plants have to be constructed and placed near the airports.
Such plants will require regular inputs of feedstock of Jatropha. Hence, the transportation of
Jatropha will be done to these plants near the airport. Jatropha comes in barrels which can be placed in any transportation vehicles. Since, the fuel shows many similarities with the existing
Jet A fuel chemically, the existing transportation systems and methods can be used with minimal changes. The existing Jet fuel trucks should be compatible with the new fuel along with the fuel feeding systems into the plane. Storing the Bio SPK fuel can be done in the existing storing facilities with the minimal changes of controlling the sunlight exposure and storage temperature.
67 Due to the similar nature of the Bio SPK fuels to the Jet A fuel, this is a good choice in terms of storing, implementing and transporting it to the airports.
11.5 New Advanced Systems
The newly introduced and advance technology that requires more detailed outlook are explained in this section. These include NextGen, Lidar/Optical Sensing Interface and more.
11.5.1 NextGen
NextGen is an initiative by the Federal Aviation Administration (FAA) to transition from a ground, radar based surveillance system, to a satellite, GPS system. The goals of this initiative are to reduce air traffic congestion in the air and at airports while improving safety and emissions. Many reductions in delays and emissions will come from the optimization of approach and departure paths. This will allow the airport to prepare departures and arrivals depending on the air traffic around the airport. Also, this will allow for aircraft to operate in the most fuel efficient manner possible, greatly reducing delays. This will be aided by improved wake vortex models, which will allow a better prediction of when wake vortices have dissipated.
This has the potential to increase the capacity of an airport significantly.
While the aircraft are en route to their destination, they will no longer have to travel between ground points for navigation. The use of GPS will allow the aircraft to fly directly between airports and allow for improved deviations due to weather or malfunctions. This decreases fuel burn since the aircraft can fly a shorter route to its destination. The vertical separation between aircraft will be reduced at high altitudes to potentially double the amount of aircraft that can fly in a given airspace. This allows for improved routing and fewer deviations due to surrounding traffic. A major benefit of a satellite based surveillance system is the ability to cover areas that radar cannot reach. Examples of such areas are Alaska and the Gulf of
68 Mexico, where there is large amounts of traffic. These areas will be able to have reduced separation while receiving guidance from an air traffic controller, greatly improving safety. A final major improvement is the capability to land simultaneously on closely spaced parallel runways. This is a major advancement that can greatly improve an airports capacity.
11.5.2 Lidar/Optical sensing Interface
Lidar/Optical sensing interface is a strong tool in acquiring data of the upcoming atmospheric conditions. Its Atmospheric Remote Sensing allows the pilots to efficiently understand what kind of atmospheric conditions are they about to head into well ahead in time.
The turbulence detection allows them to avoid the areas where the turbulence is high and fatal. It can detect the water content, aerosol distribution in the upcoming atmospheric area, and contrail as well.
This system beams optical rays that allow it to acquire the necessary data and create 3 D imaging for the pilots to better understand the conditions and find a safer route to the destination.
Such information is also important for study and research the atmospheric conditions for any other purposes as well.
11.5.3 GPS Landing
GPS landing is a new landing aid that will replace the existing Instrument Landing systems that are used at present. Such aid will be very helpful in landing aircrafts under serious and dangerous climate conditions where visibility is reduced to near zero.
The GPS landing system calculates the three dimensional position of the incoming aircraft for high accuracy. The calculations are compared to the runway’s position and elevation.
Using these, the pilot is able to land the aircraft by steering and reviewing on a display inside the cockpit. This way, Autoland can also be used which can make landings more efficient, faster
69 and smoother. Autoland is a technology that allows the on board computer to land the airplane using the necessary information required for landing from the GPS.
Technologies such as the GPS landing, Lidar optical sensing and NextGen can make air travel safer, smoother and more efficient. Through efficiency, costs cut can also be expected.
12. Cost Estimation
The aircraft cost estimation is a tedious process which includes all aspects starting from the design and manufacturing of the aircraft to the operating and maintenance costs of the proposed aircraft model. Based on the RFP requirements, the drivers that influence the aircraft direct operating cost are as follows:
• Increase L/D – Increase the aerodynamic and structural efficiency of the aircraft
• Specific Fuel Consumption – Research and analyze the best alternative fuel for the
aircraft
• Take off gross weight – Increase the aerodynamic, propulsion and structural efficiency
Such drivers, as mentioned above can affect the fuel use of the aircraft. Based on Roskam’s approach to estimate all the costing aspects of the proposed aircraft model, the cost estimation is divided into three main divisions:
1. Cost of Research, Development, Testing and Evaluation
2. Cost of Manufacturing (Both lifetime and unit cost)
3. Operating Cost of the aircraft
70 12.1 Cost of Research, Development, Testing and Evaluation
In order to compute this cost, the take off gross weight is computed preliminarily. The take off gross weight is computed to be 138,797.66 lb. Working through Roskam’s approach for cost estimation, the costs of the following areas were computed:
1. Cost of Airframe and Development
2. Cost of development and support
3. Cost of manufacturing
4. Cost of materials
5. Cost of tooling
6. Cost of quality control
7. Cost of flight tests operation
8. Cost of engines and avionics
The cost of engines and avionics is assumed to be $15,000,000 since there is not enough
information available to acquire the cost of the engine selected. The factor of computer aided
designing and electronic designing is taken to be at the highest level of 0.8 since the use of
computer aided designing has been performed intensely. The research involved for this design
has been highly dependent on the use of computer models, computer analysis and user interface
designs. Also, the strut braced design requires more designing than the regular conventional
aircrafts which can be easily facilitated by computer designing software and interface. Hence,
keeping these factors in mind, the factor of the computer aided design has been chosen higher
than the existing CAD factor for conventional aircrafts. The difficulty factor is taken at a
medium level of 1.5. This number is selected by the team since the aircraft structure shares some
71 similarities with the conventional design of the present aircrafts. The computations and values for the cost of research, development, testing and evaluation are shown in the table 12.1.1 below:
Table 12.1.1 Cost of Research, Development, Testing and Evaluation
W(Take off) lb 138797.666 W(ampr) lb 43570.1826 V(max) ft/s 478 N(RDTE) 4 Fdiff 1.5 Fcad 0.8 Man Hours 3511229.206 C(airframe & Develop) $ 228,229,898.37 C(development support) $ 78,711,930.90 Man Hours(manufacturing) 6944780.194 C(manufacturing) $277,791,207.76 C(materials) $74,988,697.96 C(tooling) $100,417,085.26 C(quality control) $36,112,857.01 C(flight tests operation) $15,438,590.42 C(engines &Avionics) $15,000,000.00 C(RDTE) $1,033,000,000.00
The final cost of research, development, testing and evaluation is computed to be
$1,033,000,000.00. This cost value matches well with the similar cost estimations of other aircrafts in its size sector/level. [21]
12.2 Cost of Manufacturing – Lifetime and Unit Cost
The cost of manufacturing is an important part of the total cost estimation process. It allows the buyer to have a future estimation of the unit price and capital requirement to manufacture the strut braced design. The cost of manufacturing has the following sub divisions:
1. Cost of airframe engineering and development
2. Cost of airplane program production
72 3. Flight test operation cost
In order to compute the total manufacturing cost, the total numbers of airplanes produced in the airplane history is taken to be 2500. This is based on an average number of airplanes produced for the specific design in their lifetime. The total numbers of airplanes manufactured per month is taken to be 2. With inclusion of the man hours required to produce the plane and the salaries involved, the following computation of the total manufacturing cost of the airplane design was computed. Table 12.2.1 shows the step by step computation of the total manufacturing cost:
Table 12.2.1 Total manufacturing cost of the strut braced airplane design
Total Number of engineering manhours MHR (aed) 11405260.87 required for the entire airplane program $ Cost of airframe engineering and design C(aed) man. 227,980,536.59 manufacturing C(engine+avion.) 18750000000 Cost(interior) 1875000000 Cost(man.) $7,815,027,857.83 Labor Cost incurred in Manufacturing. MHR(man.) 202320476.6 Manhours manufacturing Total material cost associated with building Cost(material) prog 12299364856 Np airplanes Cost(material) man $12,224,376,157.71 Materials costs incurred while manufacturing total number of tooling manhours to prepare MHR(tooling) 22835088.43 2504 planes Cost(tooling) man. $812,986,451.76 Tooling Cost Cost(quality Control) 1015953622 Quality Control Cost C(apc) man. $42,493,344,088.81 Airplane Program Production Cost Cost(operating cost) Flight Test Hours 2 Overhead Factor 4 C(Fto) 0 Flight Test Operation Cost C(finance) Cost of Financing the Program C(Profit) The Profit made based on 10 percent C(Manufacturing) $142,404,000,000.00 The Total cost of manufacturing the airplane
73 Based on the computation, the total cost of manufacturing is $142,404,415,418.02.
Hence, the unit price is acquired by dividing the total cost by 2500 (the number of total airplanes produced for this airplane design for lifetime). The unit cost is calculated to be $56,961,766.17.
This is a reasonable cost amount when compared to the other airplane designs in its segment.
[21]
12.3 Direct Operating Cost
The direct operating cost is an important aspect of the cost estimation. It covers all the aspects of the costs that are related to the direct operation of the aircraft including fuel costs, insurance, crew costs etc.
It should be noted that the direct operating costs are computed on the basis of per nautical miles. The salaries of the captain and co pilots are taken on the basis of the average salaries for the year 2009. After intense research, the average fuel price per gallon is acquired to be $1.67.
This is much lesser than the existing Jet A fuel price which is hovering around the $4.50 price.
The storage, implementation and transportation of the new fuel are very similar to the existing fuel systems. Hence, no extra cost is required in modifying the fuel systems to be compatible with the new fuel. The density and the weight of the oil is also computed through research and examples used by existing aircraft design models. Table 12.3.1 shows the direct operating cost of flight:
74 Table 12.3.1 Direct Operating Cost of Flight
t(gm) 0.19578681 Time spent in ground maneuvers t(cl) 0.25 time to climb to cruise altitude Distance to climb and to accelerate to cruise R(cl) 85 speed R(de) 116 Descent Distance R(man) 44.71173159 Distance Covered while maneuvering T(cruise) 4.060198124 cruise time t(block) 5.105984934 U(ann) 17624.79639 Annual utilization in block hours Vbl 394.8307772 Kj 0.26 Factor of Crew Cost Salary(Captain) 150,420 Salary(Co Pilot) 88,799 AH 800 Number of Flight House TEF 18.56 travel Expense per block hour Crew Cost(Captain) 0.647040491 per nautical mile Crew Cost(Co pilot) 0.401231196 per nautical mile Total Crew Cost 1.048271687 per nautical mile Cost(pol) 1.5995738 Fuel and oil cost per nautical mile W(block fuel) 13664 fuel Price $1.67 Fuel Density 7.43 Weight(ol) 7.148378908 Weight of oil and lubricants Oil Density 7.74 f(insurace) 0.08 Annual hull insurance rate AMP 98680000 Market Price Cost(Insurance) 1.134446502 per nautical mile DOC(flight) 3.782291989 Direct operating Cost of Flight
The direct operating cost of maintenance is computed based on the following drivers:
1. Cost of labor for airframe and systems
2. Cost of labor for engine
3. Cost of materials for airframe
4. Cost of materials for engine
75 The direct operating cost of maintenance is computed step by step and shown in Table 12.3.2 as follows:
Table 12.3.2 Direct Operating Cost of Maintenance
MHR(map bl) 12.5 Number of Airframe and systems maint. Rlap 20.11 Airplane Maintenance labor rate Cost(labor) ap 0.655765115 Airframe and Systems MHR(meng) 2.5 Number of Engine Maint. Hours R(l eng) 21.83 Engine Main. Labor rate Cost(labor engine) 0.370163266 C(material apblhr) 380 Airframe and systems maintenance C(mat ap) 0.991310765 C(mat/engblhr) 170 engine maintenaince material cost Cost(materials engine) 1.153050943 Cost of applied maintenance burden per Cost(amb) 2.631499645 nautical mile DOC(Maintainence) 5.801789734 Direct Operating Cost of Maintenance
The direct operating cost of depreciation per nautical mile is computed using the following drivers:
1. Cost of airplane depreciation w/o engines, avionics and spares
2. Cost of engine depreciation per nautical mile
3. Cost of depreciation of avionics systems
4. Cost of engine spare parts depreciation
Using Roskam’s methods and formulae, the direct operating cost of depreciation per nautical mile is computed in Table 6.3.3 and shown below:
76 Table 12.3.3 Direct Operating Cost of Depreciation per Nautical Mile
Cost of Airplane Depreciation w/o engines, C(dap) $ 0.54236 avionics and spares f(dap) 0.85 airplane depreciation factor AEP $ 56,961,766.17 Airplane Estimated Price N(e) 2 Number of Engines Per Plane EP $ 6,000,000.00 Engine Price ASP $ 5,000,000.00 Avionics System Price DP(ap) $ 9.00 Airplane Depreciation Period U(annual bl) 17624.79639 Annual Utilization in block hours V(bl) 394.8307772 Block speed in nm/hr F(deng) 0.85 Engine Depreciation Factor DP(eng) 7.00 Engine Depreciation Period C(deng) $ 0.20940 Cost of engine depreciation per nautical mile C(dav) $ 0.14370 Cost of depreciation of avionics systems F(dapsp) 0.85 airplane spare part depreciation factor F(apsp) 0.1 airplane spare parts factor DP(apsp) 10 Airplane spare parts depreciation period C(dapsp) $ 0.05492 Cost of depreciation of airplane spare parts F(dengsp) 0.85 Engines Spare Parts depreciation factor F(engsp) 0.5 Engine spare Parts factor ESPPF 1.5 Engine spare parts price factor DP(engsp) 7 Depreciation period for engine parts C(dengsp) $ 0.15705 Cost of engine spare parts depreciation Direct operating cost of depreciation per DOC(depr) $ 1.11 nautical mile
The direct operating cost of landing fees and registry taxes is computed per nautical mile as well.
Table 12.3.4 shows the direct operating cost of landing fees and registry taxes:
Table 12.3.4 Direct Operating Cost of Landing Fees and Registry Taxes
C(aplf) 277.595332 Airplane landing fee per landing C(lf) 0.137696097 Landing fees F(rt) 0.002387977 Factor of registry taxes C(rt) $ 0.03 Cost of registry taxes per nautical mile Direct operating cost of landing fees and DOC(lnr) $ 0.16522 registry taxes
77 Hence, the direct operating cost of landing fees and registry taxes comes out to be $0.16. The
direct operating cost of financing the airplane is computed by multiplying the interest rate for the
given year with the total direct operating cost. Table 12.3.5 shows the direct operating cost of
financing the airplane per nautical mile.
Table 12.3.5 Direct Operating Cost of Financing the Airplane
Direct operating cost of financing the airplane per DOC(fin) $ 0.80692 nautical mile
Hence, by adding all the direct operating costs computed and shown above, the total
direct operating cost is acquired. The total direct operating cost is summed up to be:
Table 12.3.6 Total Direct Operating Cost
Total DOC $ 11.53 Total Direct Operating Cost per nautical mile
The total direct operating cost is $11.53 per nautical mile, which is a good and acceptable
value for operating the strut braced airplane design. With 175 passengers, the “per passenger, per
nautical mile” operating cost is 6.4 cents. This is a reasonable value when compared to the
existing aircraft cost analysis. According to USA Today, the cost of flying one passenger for one
nautical mile hovers from 6.8 cents to 21.3 cents [31]. Also, the 6.8 cents operating cost is for
short haul flights using smaller airplanes and budget airlines. Hence, the selected design is a very
good choice in terms of cutting costs for long haul flights. [21]
78 13. Concluding Remarks
As justified in the preceding analysis, the Environautics EN 1 meets or exceeds all requirements put forward by the AIAA Undergraduate Design Competition Request For
Proposal. The strut braced design of the aircraft gives it a significant advantage over the currently used conventional transports, such as the Boeing 737 and the Airbus 320. The strut permits an increase in wingspan and aspect ratio, and decrease in wing thickness and sweep angle, that allows for a more aerodynamically efficient aircraft. Composite materials help to drive down the weight of the aircraft by providing strong, light materials that can withstand the stresses imposed on the structure during flight. The new engine technology burns less fuel per pound of thrust and generates less noise than engines currently in use. The gains in aerodynamic, structural and propulsive efficiency allow for a drop in required fuel burn. Combined with the more environmentally friendly SPK biofuel, the environmental footprint of the EN 1 is reduced dramatically. Finally, the cost to operate the EN 1 is considerably less than the Boeing 737 and
Airbus 320, making it an attractive choice for new aircraft purchases by the year 2020.
79 Table 13.1 RFP Compliance Summary
Design Factor Requirement EN 1 Compliance Safety and Airworthiness FARs FARs Yes Regulations Crew 2 2 Yes Passengers 175 (1 Class) 180 (1 Class) Yes Seating Pitch 32”, Width 17.2” Pitch 32”, Width 17.4” Yes Width > 12.5 ft Width = 13.28 ft Cabin Dimensions Yes Height > 7.25 ft Height = 8.335 ft Cargo Volume 1,240 ft 3 1,240 ft 3 Yes Takeoff Distance 8,200 ft 4,388 ft Yes Landing Speed < 140 KCAS 135 KCAS Yes > 95,000 lbs Maximum Landing (Maximum zero fuel weight plus 115,000 lbs Yes Weight fuel reserves for maximum range) Cruise: Mach 0.8 Cruise: Mach 0.8 Operating Speed Yes Maximum: Mach 0.83/340 KCAS Maximum: Mach 0.83 Initial: 35,000 ft Initial: 35,000 ft Cruise Altitude Yes Maximum: 41,000 ft Maximum: 41,000 ft Nominal: 1,200 nm Nominal: 1,200 nm Range Yes Maximum: 3,500 nm Maximum: 3,900 nm Payload Capability 37,000 lb 39,860 lb Yes
80 12. References
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[21] Roskam, Jan. Airplane Cost Estimation Design, Development, Manufacturing and Operating: Airplane Design Part VIII . Darcorporation, 2002.
[22] Penner, Joyce E., “Aviation and the Global Atmosphere,” IPCC Special Reports on Climate Change , ch.7.
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