An international cooperative design effort between Virginia Tech and Loughborough University presents:
May 12, 2009
JFTL – C-328 Ostrich Final Report
Executive Summary
The Virginia Tech and Loughborough University International Aircraft Design team’s C-328 Ostrich fulfills an existing need in the United States military and its allies for an Extreme Short Take Off and Landing (ESTOL), transonic cruise aircraft. The requirements originate from within the U.S. Army Joint Heavy Lift (JHL) program, the U.S. Air Force Advanced Joint Air Combat System (AJACS), and a Special Forces mission originating from the Iran Hostage Crisis. Due to budget reductions, these programs were harmonized into the single Joint Future Theater Lift (JFTL) program in pursuit of a multi-role tactical transport capable of operating at hot and high field conditions. The JFTL mission requires a Mach 0.8 cruise with either a 328 ft takeoff, 26,000 lb payload and 1000 nm combat radius or 1,500 ft takeoff, 66,000 lb payload and a 500 nm combat radius. The C-328 aircraft employs an innovative Distributed Propulsion system in conjunction with blown flaps and 2 large under wing turbofans as the solution to this challenge. All requirements were met, except the 328 ft take off and landing at hot and high conditions, where 571 ft is required for ESTOL. In response to the JFTL program, the Virginia Tech and Loughborough University International Aircraft Design team present the C-328 Ostrich.
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65.21 50.55 SECTION A-A CLEAN CONFIGURATION 2.89 26.31 20.54 VIEW SHOWING TAKEOFF TAIL SECTION A-A DISTRIBUTED PROPULSION TAKEOFF CONFIGURATION 10.17 CONFIGURATION LANDING GEAR LAYOUT (REF)
SECTION B-B DATA SUMMARY CLEAN CONFIGURATION
PERFORMANCE 9.86
MAX RANGE AT 12T PAYLOAD 4917 nm MAX RANGE AT 30T PAYLOAD 3289 nm MAX FERRY RANGE 5510 nm 2.51 CRUISE MACH NUMBER 0.8 SECTION B-B CRUISE ALTITUDE 35,000 ft 97.77 177.49 DISTRIBUTED PROPULSION TAKEOFF CONFIGURATION 110.14 POWER PLANT
MAIN ENGINES 2 X ROLLS-ROYCE TRENT 895 B B DISTRIBUTED ENGINES 36 X HONDA/GE HF-120 TRENT 895 THRUST 93,400 lb HONDA/GE HF-120 2,050 lb SECTION B-B C C SPOILER UP AND FLAP DOWN CONFIGURATION WEIGHTS MAC POSITION OPERATIONAL WEIGHT EMPTY 143,141 lb MAX TAKE OFF WEIGHT 330,693 lb A A DIMENSIONS
LENGTH 147.9 ft WING SPAN 177.4 ft SECTION C-C WING AREA •••••••••••••••••••• CLEAN CONFIGURATION ASPECT RATIO 4.97 WING LEADING EDGE SWEEP •••••••••••••••••• • ••••••••••••••••••••• WHEEL BASE 40.38 ft WHEEL TRACK 26.31 ft TAIL SCRAPE ANGLE •••••••••••••••••• • ••••••••••••••••••••••• 4.85
SECTION C-C DISTRIBUTED PROPULSION TAKEOFF CONFIGURATION
JFTL 61.62 24.26
27.15 KEY 81.08 148.21 DISTRIBUTED PROPULSION ENGINE 1.83 INTERNATIONAL DESIGN TEAM
VT MEMBERS LU MEMBERS
CONTROL SURFACE Tyler Aarons David Brindley David Gladson Victoria Cope Ryan Meritt Scott Ferry Chris Olien Ryan Hurril Ben King Grant Parrish CENTER OF GRAVITY Simon Langley VIEW LOOKING DOWN Wendy Pifer Alex McMillan Steve Sikorski Chris Skinner SHOWING FUSELAGE Shadie Tanious Seb Wilkes METERS 0 510152025 FRAME LOCATIONS SCALE 1:150 9 0 20 40 60 80 ISSUE FEET JFTL GENERAL ARRANGEMENT DRAWING 10-APRIL-09 UNITS FEET
JFTL – C-328 Ostrich Final Report
Table of Contents
1. The JFTL Team...... 14 1.1 History of the Collaboration...... 14 1.2 New Perspective ...... 14 1.3 The 2008-2009 Virginia Tech International Design Team ...... 14 2. Introduction ...... 16 2.1 Combination of Existing Program Requirements ...... 16 2.2 Final Aircraft Performance Requirements...... 17 3. Design Evolution ...... 19 3.1 Early Design ...... 19 3.2 Joint Concepts ...... 19 3.3 Concept Downselection ...... 19 4. Initial Sizing ...... 21 4.1 Initial Concept Sketching ...... 21 4.1.1 Wing Geometry ...... 21 4.1.2 Fuel Burn ...... 21 4.1.3 Center of Gravity...... 22 5. Aerodynamics ...... 23 5.1 Airfoil ...... 23 5.2 Wing...... 25 5.3 L/D Optimization ...... 27 5.4 Flaps ...... 27 5.5 Inlets ...... 28 6. Propulsion ...... 30 6.1 Distributed Propulsion System ...... 30 6.1.1 Principle...... 30 6.2 Jet Flap Theory and Integration ...... 31 6.2.1 Principle...... 31 6.2.2 Sizing ...... 32 6.2.3 Stage Flap Deflection ...... 33 6.3 Distributed Propulsion Design ...... 34
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6.3.1 Installation ...... 34 6.3.2 Number of Engines ...... 34 6.3.3 Final Engine Selection ...... 35 6.3.4 Performance ...... 37 6.4 Cruise Engine System ...... 38 6.4.1 Engine Selection...... 38 6.4.2 Nacelle Design...... 40 6.4.3 Reverse Thrust ...... 41 6.4.4 Positioning ...... 42 6.4.5 Performance ...... 44 7. Structures...... 46 7.1 Velocity-Load Diagram ...... 46 7.2 Wing Box Layout ...... 47 7.2.1 Spars and Ribs ...... 48 7.2.2 Integration of Distributed Propulsion ...... 50 7.2.3 Flap Attachment ...... 51 7.2.4 Finite Element Model ...... 51 7.3 Other Structural Components ...... 52 7.3.1 Fuselage ...... 52 7.3.2 Horizontal and Vertical Stabilizers ...... 53 7.4 Materials ...... 53 8. Weight and Balance...... 55 8.1 Empirical and Group Methods Used ...... 55 8.2 Center of Gravity...... 57 9. Stability and Control ...... 59 9.1 Tail Sizing ...... 59 9.2 Control Surface Sizing ...... 60 9.3 Longitudinal Stability ...... 60 9.3.1 Trimming for Takeoff ...... 60 9.3.2 Dynamic Longitudinal Stability ...... 61 9.4 Lateral-Directional Stability...... 62 10. Systems ...... 66 10.1 Landing Gear ...... 66 Senior Design Project 8 May 2009
JFTL – C-328 Ostrich Final Report
10.1.1 Layout and Arrangement ...... 66 10.1.2 Tires ...... 68 10.1.3 Main Gear Housing and Structure ...... 69 10.1.4 Special Features ...... 70 10.1.5 Pilot Control and Operation ...... 71 10.2 Fuel Systems ...... 72 11. Performance ...... 73 11.1 Powered Lift Mission Segments ...... 73 11.1.1 Short Takeoff Ground Roll ...... 74 11.1.2 Short Landing Ground Roll ...... 77 11.2 Conventional Mission Segments ...... 79 11.2.1 Unrestricted Takeoff and Landing ...... 79 11.2.2 Climb, Cruise and Descent ...... 79 11.2.3 Loiter and Idle Segments ...... 80 11.3 Mission Simulation ...... 80 11.4 Range and Endurance ...... 81 12. Cost...... 83 12.1 Introduction to Aircraft Associated Costs ...... 83 12.2 Current Military Transport Market ...... 84 12.3 Estimating RDT&E, Flyaway, and Unit Cost ...... 86 12.4 Estimating Operation, Maintenance, and Disposal Costs...... 88 12.5 Life Cycle Costs ...... 90 13. Conclusion ...... 91
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List of Figures
Figure 2.2.1 - Special Forces Mission Schematic ...... 17 Figure 2.2.2 – AJACS Mission Schematic ...... 17 Figure 3.3.1 - Design Evolution ...... 20 Figure 4.1.1 - Initial Concept Sketching spreadsheet ...... 22 Figure 5.1.1 - Sketch Demonstrating a Typical Supercritical Airfoil Shape ...... 24 Figure 5.1.2 - XFOIL Analysis of Whitcomb Airfoil ...... 25 Figure 5.2.1 - Wing Configuration from Tornado ...... 25 Figure 5.2.2 - Delta Cp Distributions ...... 26 Figure 5.2.3 - Lift Distribution Across the wing ...... 26 Figure 5.4.1 - Cross Sectional View of the Blown Flap System...... 27 Figure 5.5.1 - Example of a NACA Inlet ...... 28 Figure 6.1.1 - Distribute Propulsion Design ...... 30 Figure 6.1.2 - DP Impact on Lift Coefficient Distribution...... 31 Figure 6.2.1 - Diagram of a Jet-Flapped Airfoil ...... 32 Figure 6.2.2 - Blown Span/ Jet-Flapped Wing ...... 33 Figure 6.2.3 - Net Forward Thrust from Distributed Propulsion at Various Flap Angles ...... 34 Figure 6.3.1 - Engine Size Limitation ...... 35 Figure 6.3.2 - Engine Characteristics ...... 36 Figure 6.4.1- Two versus Four Engine Design Illustration ...... 39 Figure 6.4.2 - Rolls Royce Trent 895 Engine ...... 40 Figure 6.4.3- Long Duct Nacelle ...... 41 Figure 6.4.4 - Nacelle Dimension Nomenclature ...... 41 Figure 6.4.5 - Reverse Thrust Comparison ...... 42 Figure 6.4.6 - Horizontal Engine Placement ...... 43 Figure 6.4.7 - Forward Engine Placement ...... 43 Figure 6.4.8 - Trent 895 SFC Thrust/Altitude Curves ...... 44 Figure 6.4.9 - Trent 895 SFC Thrust Curves ...... 45 Figure 7.1.1 - V-n Diagram ...... 47 Figure 7.2.1 - Wing Box Structure showing Distributed Propulsion Engines ...... 48 Figure 7.2.2 - Wing Box Layout ...... 49 Figure 7.2.3 - Cross-Sectional Diagram of Wing at Root ...... 50
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Figure 7.2.4 – Cross-Sectional Diagram of Wing at Tip of Distributed Propulsion Section ...... 51 Figure 7.2.5 – Meshed Finite Element Model of Wing Box ...... 52 Figure 7.3.1 – Top Down View of Fuselage showing Frame Locations...... 53 Figure 7.4.1 – Material Weight Breakdown ...... 54 Figure 8.2.1 – Side View with CG Location...... 57 Figure 8.2.2 - Potato Plot showing the CG Envelope of the AJACS Mission...... 58 Figure 8.2.3 Potato Plot showing the CG envelope of the Special Forces Mission ...... 58 Figure 10.1.1 - A400M Landing Gear ...... 66 Figure 10.1.2 - Tri-twin Tandem Landing Gear Arrangement ...... 67 Figure 10.1.3 – Aft towing angle ...... 68 Figure 10.1.4 – Tail Tipping Angle ...... 68 Figure 10.1.5 – Individual Tire Loading ...... 69 Figure 10.1.6 - Sponson Configuration ...... 70 Figure 11.1.1 - Free Body Diagram During ESTOL Ground Roll...... 73 Figure 11.1.2 – Flap Deflection Schedule for Hot and High Conditions...... 75 Figure 11.1.3 - Flap Deflection Schedule for Sea Level Conditions...... 76 Figure 11.1.4 – Landing Acceleration with Effect of Various Arresting Systems...... 78 Figure 11.4.1 - Payload vs. Range Diagram ...... 82 Figure 12.1.1 - Elements of Life Cycle Cost ...... 83 Figure 12.2.1 - Learning curve affect on cost: JFTL with comparable aircraft ...... 86 Figure 12.3.1 - Tabulated RDT&E + Fly-away Costs ...... 87 Figure 12.3.2 - Unit RDT&E + Fly-away Costs ...... 88 Figure 12.4.1 - Operation and Maintenance Costs ...... 89 Figure 12.4.2 - O&M Costs per year ...... 89
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List of Tables
Table 2.1.1 - Individual Program Requirements ...... 16 Table 2.2.1 - Final Aircraft Performance Specifications ...... 18 Table 6.3.1 - Distributed Propulsion Decision Matrix ...... 36 Table 6.3.2 - HF120 Engine Characteristics ...... 36 Table 6.3.3 - Summarized Distributed Propulsion Engine Performance ...... 37 Table 6.4.1 - Two Versus Four Engine Design Comparison...... 38 Table 6.4.2 - Candidate Cruise Engines ...... 39 Table 6.4.3 - Rolls Royce Trent 895 Engine Characteristics...... 40 Table 6.4.4 - Nacelle Dimensions ...... 41 Table 6.4.5 - Summarized Distributed Propulsion Performance ...... 44 Table 7.2.1 - Web Thicknesses at Root and Tip for Wing Box Spars ...... 49 Table 8.1.1 - Empirical Data for Approximate Empty Weight Buildup ...... 55 Table 8.1.2 – Component Group Weights and Moments ...... 56 Table 9.1.1 - Summary of Tail Surface Sizing...... 60 Table 9.2.1 – Summary of Control Surface Sizing ...... 60 Table 9.3.1 – Mass Moments of Inertia ...... 61 Table 9.4.1 - Lateral Directional Stability and Control Derivatives ...... 63 Table 9.4.2 - Spiral Mode Minimum Allowable Time to Double Amplitude, ...... 63
Table 9.4.3 - Du nD ...... 64
Table 9.4.4 - Dutchtch Roll UndampedDamping Ratio, Natural ...... Frequency, ω ...... 64 Table 9.4.5 - Dutch Roll Real Root Part Value, ...... 64 Table 9.4.6 - Roll Mode Maximum Allowable Tim Constant, ...... 65 Table 10.2.1 - Fuel Tank Capacity Specifications ...... 72 Table 11.3.1 - Mission Analysis Summary ...... 81 Table 12.2.1 - Performance measurements and number of comparable aircraft ...... 85 Table 12.5.1 - Estimated JFTL Life-Cycle Costs of 200 Aircraft over 30 years service life ...... 90
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Nomenclature
English Acronyms AC – Aerodynamic center ACS – AirCraft Synthesis AOA – Angle of attack AJACS - Advanced Joint Air Combat System AR – Aspect ratio Program b - Span ANSYS – AirCraft Synthesis c - Chord AOE – Aerospace and Ocean Engineering c = Wing mean chord AVID – Air Vehicle Integrated Design c/cf – Wing chord to flap chord ratio CER – Cost Estimating Relationship - Lift Coefficient CFD – Computational Fluid Dynamics CG – Center of gravity CVO – Chief Visionary Officer cht = Horizontal tail volume coefficients DAPCA – Development and Procurement Cost of cvt = Vertical tail volume coefficient Aircraft DP – Distributed propulsion DARPA – Defense Advanced Research Projects F – Lift correction for aspect ratio Agency Kg- Gust correction factor DoD – Department of Defense L - Lift ESTOL – Extreme Short Takeoff and Landing L/D – Lift to drag ratio FAR – Federal Aviation Regulations Lht = Horizontal tail moment arm FEM – Finite Element Model Lvt = Vertical tail moment arm FOD – Foreign Object Debris M – Mach Number ICS – Initial Concept Sketching mjvj – Jet-momentum JFT – Jet Flap Theory S – Planform Area JFTL – Joint Future Theatre Lift Program S’ – Wing blown area JHL – Joint Heavy Lift Program sfc – Specific fuel consumption LU – Loughborough University Sht = Horizontal tail area MAC – Mean Aerodynamic Chord Svt = Vertical tail area NACA – National Advisory Committee on – Planform exposed area Aeronautics – Planform wetter area O&M – Operations and Maintenance t – Thickness OEI – One Engine Inoperative t/c – Thickness to chord ratio OWE – Operating Weight Empty T/W – Thrust to Weight ratio RDT&E – Research, Development, Testing, and TOGW – Takeoff Gross Weight Evaluation V - Velocity STOL – Short Takeoff and Landing STP – Standard Temperature and Pressure Greek USAF – United States Air Force USB – Upper Surface Blowing - Blowing coefficient VT – Virginia Tech Cl/∂δ –Lift curve slope VTOL – Vertical Takeoff and Landing α – Airfoil incidence β∂ = Sideslip angle δ – Flap deflection angle Units λ – Part span of jet flaps °F = Fahrenheit Λ - Sweep ft = Feet ν – Fuselage cut-out area hr = Hour ρ - Density in = Inches kts = Knots lb = Pounds sec = seconds
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JFTL – C-328 Ostrich Final Report
1. THE JFTL TEAM
The collaboration responsible for the design of the C-328 Ostrich is a group of students who had the unique opportunity to work together across the Atlantic Ocean. The team has been established throughout the years as a special connection between two universities as well as a foundation for future relationships as budding professionals.
1.1 History of the Collaboration The VT/LU design collaboration was started twelve years ago by Dr. Jim Marchman (VT) and Dr. Gary Page (LU). The Virginia Tech Aerospace and Ocean Engineering (AOE) department has sponsored this group along with other several other funding organizations in order to foster the spirit of design amongst students with different technical backgrounds and cultural perspectives.
After serving for many years as advisor and facilitator for the project, Dr. Marchman decided to phase the project out as he prepared for retirement. Upon discovery that the collaboration would be coming to an end, Sam Wilson, III, the Chief Visionary Officer (CVO) of AVID Aerospace volunteered to the VT AOE department to take the project on as the VT advisor.
1.2 New Perspective Naturally, the sort of “change in command” has caused the project to take on a new angle. Despite some difficulties in initial logistics (funding, travel plans, accommodations, etc.) the VT team was still successful in maintaining the tradition of the program by travelling to England in the fall and hosting the British team in the spring.
Naturally, the design problem for the year was also of a different sort. Because of its complexity, more emphasis was placed on the design process and understanding of the trades involved in aircraft design than the final detail of the design, as was the case in the past.
1.3 The 2008-2009 Virginia Tech International Design Team This year’s team consists of all Virginia Tech AOE seniors, concentrating in aircraft design. The students represent a good sample of the senior class, with members of various backgrounds and interests regarding the subject of aircraft design.
The members of the team and their positions are shown below:
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Tyler Aarons – Mission/Performance and David Gladson – Structures Report Coordinator
Ryan Meritt – Propulsion Chris Olien – Aerodynamics and Configuration/CAD
Grant Parrish – Weights Wendy Pifer – Stability and Control
Steve Sikorski – Cost/Economics and Shadie Tanious – Team Leader and Systems Configuration/CAD
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2. INTRODUCTION
2.1 Combination of Existing Program Requirements There currently exists a need for a tactical transport, short takeoff and landing (STOL) aircraft for the United States armed forces. This need served as the launching platform for the U.S. Army based Joint Heavy Lift (JHL) program and a U.S. Air Force Advanced Joint Air Combat System (AJACS) program. Both of these programs require a heavy lift vehicle but have additional, conflicting mission requirements which drove them to run independently. The JHL program requirement was for a vertical takeoff and landing (VTOL) vehicle capable of cruising at speeds in excess of conventional rotorcraft, while being able to carry a payload of 44,092-57,320 lbs. In contrast, the AJACS program placed emphasis on replacing the aging C-130 fleet with an aircraft capable of a 1,500ft STOL operation and a higher cruise speed capability of Mach 0.8, with a larger 66,138 lbs payload.
Ideally, these programs would have produced two separate aircraft which would be able to satisfy their individual specifications; however, funding constraints have forced the U.S. Army and U.S Air Force to merge their separate pursuits of a future tactical transport. This has created a new program called the Joint Future Theatre Lift (JFTL) program. Funded by the Air Force, the JFTL aims to develop an aircraft capable of satisfying both of these missions with a single vehicle. In addition, the U.S. Special Forces required a vehicle with extremely short takeoff and landing (ESTOL) capability to deploy troops and equipment in hostile environments. These Special Forces requirements were also fed into the JFTL program. The above requirements are tabulated below in Table 2.1.1.
Table 2.1.1 – Individual Program Requirements
JHL (Army) AJACS (Air Force) Special Forces
Takeoff Run: 0 1500 ft 328 ft
57,300 lb 66,200 lb 26,400 lb Payload: - Stryker - 7 x 463 L pallets - 2 HMMWVs + 12 man crew
Cruise Speed: Mach 0.4 Mach 0.8 Mach 0.8
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In order to best suit the needs of the aforementioned customers, the Virginia Tech/ Loughborough University design team derived a list of requirements for the unified JFTL design. Because the JFTL program is funded by the Air Force, the requirements of the AJACS program were given highest weight when compared to those of the JHL and Special Forces missions. The following requirements therefore represent the best compromise among the three sets specified above by the AJACS, JHL, and U.S. Special Forces missions.
2.2 Final Aircraft Performance Requirements The final compromise called for a STOL aircraft capable of carrying out both the AJACS and Special Forces missions. Figure 2.2.1 and Figure 2.2.2 show mission schematics of both the Special Forces and AJACS missions. Additionally, Table 2.2.1 below indicates the final set of performance requirements set out for the JFTL aircraft.
1000nm cruise, M=0.8, 35,000 ft
SPECIAL FORCES 26,500 lb payload
Unrestricted TO / LD 328 ft ground roll 1.5 hour idle
Figure 2.2.1 - Special Forces Mission Schematic
500nm cruise, M=0.8, 35,000 ft 45 minute loiter
AJACS 66,200 lb payload
Unrestricted TO / LD 1,500 ft ground roll
Figure 2.2.2 – AJACS Mission Schematic
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Table 2.2.1 - Final Aircraft Performance Specifications
Performance Specification AJACS Special Forces
Operational Take off/Landing Less than 4,000 ft Less than 4,000 ft
Cruise Mach 0.8 Mach 0.8
Cruise Altitude 30,000 ft 30,000 ft
Operational Radius 500 nm 1,000 nm
Tactical Ground Roll 1500 ft 328 ft (Takeoff and Landing)
Load/Unload time 20 min 60 min
Loiter 45 min 30 min
Operating Temperature 95° F 95° F
Mission payload 66,200 lb 26,400 lb
From these requirements, the VT/LU design team set out to create an innovative design to meet the requirements of tomorrow’s armed forces. The solution, as presented in this report, is the C-328 Ostrich – a multi-role strategic transport using highly advanced technology and sound principles of aircraft design in order to perform in the future combat theater.
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3. DESIGN EVOLUTION
3.1 Early Design In the early stages of the design process, the Virginia Tech and Loughborough teams were working on vastly different designs. A lack of proper communication led to dissimilar design requirements and consequently very different initial concepts. These concepts appear on the first row of Figure 3.3.1. The Loughborough concepts show many VTOL capable rotorcraft designs due to the inclusion of the JHL requirements from the Army59. The Virginia Tech concept shows the focus on a Special Operations ESTOL mission modeled after the Credible Sport C-130 mission during the Iranian hostage crisis58.
3.2 Joint Concepts The second row of the concept evolution diagram reveals how the two sets of requirements were merged into the JFTL mission during the joint time spent in England. Lengthy discussions in England led to the elimination of rotorcraft due to the Mach 0.8 cruise requirement, although several were still meant to be VTOL capable. Also, the introduction of a distributed propulsion system as the primary high lift technology is represented in almost every concept.
3.3 Concept Downselection In the third row, the concept pool has been narrowed to just three designs. The first is continuous wing type design with twin ducted fans. During VTOL, the ducted fans turn downwards and during cruise they rotate up to blow under the rear wing. This design also eliminates the horizontal tail since it becomes part of the rear wing. This design was cut for a variety of reasons, but mainly because it was seen as unnecessarily complex for any advantages it offered. The second design uses a novel propulsion system combining ducted fans and driving engines. The fuselage houses four turboshaft engines, each driving one of the four ducted fans on the wing tips. The advantages of keeping the heavy turbomachinery in the fuselage are the lightened wing structure and the lower inertias when rotating the engines for take off. In this design, only the lightweight fans are rotating, which greatly reduces stress on the airframe and the power required to rotate the engine. The third design was the chosen concept. It employs canards and a conventional tail to aid with rotation on take off along with distributed propulsion in the large wing. Additionally, two large turbofans are included for cruise thrust.
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JFTL – C-328 Ostrich Final Report
Row 1: Initial Concepts
Row 2: Joint Concepts
Row 4: Down Selection #2
Row 3: Down Selection #1
Row 5: Final Concept
Figure 3.3.1 - Design evolution: The top line shows the Loughborough concepts on the left and the Virginia Tech concept on the right. The 2nd line shows the 5 concepts developed in jointly after the harmonization of requirements. The bottom lines show the down sel to the final concept and the last row shows the final, optimized concept.
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4. INITIAL SIZING
After selecting a working conceptual design, the team began the initial sizing process. The entire design of the aircraft was heavily driven by the ESTOL portion of the Special Operations mission. In particular, the large wing area and distributed propulsion blown flaps were direct results of the low speed lift required for the 328 ft short take off.
4.1 Initial Concept Sketching The initial wing sizing was accomplished through an iterative method developed uniquely for this mission named Initial Concept Sketching (ICS). Beginning with an initial take off gross weight (TOGW) guess and a thrust to weight ratio (T/W) based on comparator aircraft, the maximum required acceleration was estimated. The acceleration was taken as constant and then used to find velocity after the 328 ft ground roll. This velocity was fed into the equation for the lift coefficient,
4.1 along with lift, L, approximated as TOGW, density, ρ, at “hot and high” conditions (4,000 ft altitude,
95°F), and an estimate of the required take off CL. The take off CL was estimated using the AVID report to DARPA4 and other blown flap papers45,56. Using this method, the remaining unknown in the CL equation is the wing area, S, which was returned to the user as an output.
4.1.1 Wing Geometry
The wing geometry, which includes sweep (Λ), span (b), chord (c), and thickness (t) were then selected and the maximum geometrically allowable number of distributed propulsion engines (DPE) placed in the wing. The fitting of the DPEs in the wing was based simply on comparing the thickness to chord ratio (t/c) value to the engine diameter since neither an airfoil nor an engine were chosen yet. The ICS program assumes that the optimum condition is met when the number of DPEs is maximized. The reasons for this are discussed in great detail in Section 5.1, but essentially boil down to being able to aerodynamically seal the unused DPEs during cruise.
4.1.2 Fuel Burn
Fuel burn was then calculated at each portion of the mission using specific fuel consumption (sfc) numbers listed for actual engines in the necessary thrust class. The methods
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used to estimate each portion of the mission were based on equations from Raymer39, Roskam41, and performance data from comparator aircraft such as the Boeing 747 and 737, the C-17 Globemaster III, the C-5 Galaxy, and the KC-135 Stratotanker. This gave an initial fuel estimate for the TOGW estimation. An estimated structural weight was calculated with a rubber sizing method for weight based on wing area, tail area, and fuselage size39. Finally, the payload, engine weight, structural weight, and fuel weight were summed to obtain a calculated TOGW. The input TOGW value was iterated until an error of less than 0.1 lbs with respect to the calculated value was reached to verify convergence of the ICS program.
4.1.3 Center of Gravity
The ICS method was also expanded to include center of gravity (CG) positions for all component weights such as engines, structure, payload, and fuel. These inputs were then used to calculate the overall CG. An aerodynamic center (AC) of the wing was calculated and both the CG and AC were plotted on a sketch of the aircraft. The aircraft sketch included the wing, payload, an idealized fuselage, and the position of the most outboard DP engine as shown in a selected screenshot from the ICS program shown in Figure 4.1.1.
Figure 4.1.1 - Initial Concept Sketching (ICS) method spreadsheet screenshot
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5. AERODYNAMICS
The initial task of the aerodynamics group was to estimate a lift to drag ratio (L/D), since this was a major factor for several of the other groups. The ICS method offered an initial estimate at L/D by computing a wetted aspect ratio (AR), which was then compared to empirical plots in Raymer39. The wetted AR accounted for wing and tail areas as well as an estimation of the fuselage area using cylinders and cones. The resulting L/D values of 11 to 14 were rough and used only in the initial stages of the design; however, before more accurate L/D calculations could be completed, the wing sizing needed to be refined. This was accomplished by using improved weight estimates from the weights group and a custom modeling of the short take off acceleration and flap deflection schedule from the performance group. Once complete, the wing area was frozen and work began on refining the particulars of its geometry.
The first constraint on the wing was the span. Since the landing strip is taken to be a standard soccer pitch, it was decided the wing span should be no wider than the pitch (180 ft). The second constraint was sweep, which as calculated using Equation 5.1, where the Mach number was chosen to be 0.8 for cruise.
5.1
5.1 Airfoil The next step was choosing the airfoil. Since all other comparator cargo jet aircraft use a supercritical airfoil, such as in Figure 5.1.1, it was the first type to be investigated for the C-328. It was generally agreed that the airfoil would be a supercritical variant as shown in Figure 5.1.1. For a supercritical airfoil, the maximum thickness is pushed much farther aft, near mid-chord, in order to delay and weaken the upper surface shockwave. The shockwave is caused when the air is accelerated over the top of the airfoil from a subsonic free stream velocity to a supersonic local velocity. The weaker shockwave results in improved lift and lower drag at higher Mach numbers. The supercritical shape has also shown good performance at low speed, which is important for this aircraft which will operate at extremely low speed for the ESTOL mission segments. Furthermore, the supercritical airfoil has increased internal volume due in part to its large thickness near the leading edge and also because a thicker supercritical airfoil can give the same performance as a
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thinner NACA series airfoil. Finally, the supercritical airfoil’s increased volume between mid-chord and ¾-chord is beneficial due to the placement of the DP engines.
Disadvantages of the supercritical airfoil for the Ostrich included the movement of lift further aft along the chord. Moving the lift aft increases the pitching moment and accordingly, the tail volume is increased to counter the moment. This results in a greater TOGW and drag. Having the weight of DP engines in the aft portion of the chord helps to counter the nose-down moment, but this engine placement also creates lift on the aft portion of the airfoil by blowing the flaps. In the end it was determined that the advantages of a supercritical airfoil far outweighed the disadvantages.
Figure 5.1.1 - Sketch Demonstrating a Typical Supercritical Airfoil Shape10
The Whitcomb supercritical airfoil, shown in Figure 5.1.2, was chosen because it is well tested and proven. Analysis was conducted using the XFOIL program to examine the effects of larger thickness to chord ratios (t/c) at cruise. Interest in larger t/c values was due to the need for internal volume for DP engines and fuel, as well as the improved low speed performance. The analysis indicated that a 14% maximum thickness was optimal. Ultimately, it was determined that the 14% thickness would be necessary at the tips to fit the required number of engines in the wing. However, the thickness was chosen to be 11% at the root since the large chord made a 14% thick wing unreasonable due to the required blending to the fuselage in a high-wing configuration.
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Figure 5.1.2 - XFOIL Analysis of Whitcomb Airfoil
5.2 Wing The lifting characteristics of the wing were analyzed with the MATLAB vortex lattice code Tornado40. The wing configuration input for Tornado is shown in Figure 5.2.1.
Figure 5.2.1 - Wing Configuration from Tornado
Two problems appeared from the Tornado analysis. First, the lift distribution was centered near the wing tips, and second, the majority of the lift was acting on the trailing edge of the wing. The lift distribution was moved inboard by twisting the wing tips downward. The optimal solution was achieved by dividing the wing into two sections with the inboard one containing the DPE engines and the other spanning the remainder of the wing to the tip. By adjusting the twist and then viewing the resulting spanload in Tornado, the point of maximum lift was moved within half the span. This type of spanload distribution control was necessary to prevent stall from first
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occurring at the wingtips, leading to control issues. The final inboard section twist is downward 2°, while the outboard section twist is an additional 3° downward. In addition to the twist, the wing was given an installed incidence of 3° to both provide a built-in angle of attack (AOA) during the ground roll and also to help move the lift further forward on the wing. Incidence was chosen using Tornado analysis with the goal of spreading lift more evenly between the leading and trailing edges. These results are illustrated in Figure 5.2.2. and Figure 5.2.3.
Figure 5.2.2 - Delta Cp Distributions - Unaltered wing (left), Twisted and Incidenced wing (right)
Figure 5.2.3 - Lift Distribution Across the wing - Without Twist (left), With Twist (right)
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5.3 L/D Optimization After completing the wing analysis, more detailed L/D calculations were completed using both Tornado and AirCraft Synthesis (ACS)3. ACS is an aircraft design software that AVID is actively developing and was first written in the 1970’s. As with any aircraft analysis software, ACS is a complex program requiring detailed inputs and generating lengthy outputs. To verify the results of these methods, hand calculations were used. The L/D output was generated for a range of AOA’s at several combinations of Mach number and altitude to be encountered during the mission. This gave a value of 13.4 L/Dmax.
5.4 Flaps The flaps are an integral and complex component of the design of the aircraft. Composing the trailing 30% of the wing chord and 51% of the span , the flaps are blown directly by the distributed propulsion and generate the required additional lift for ESTOL. The sizing and was accomplished by the Propulsion team under the jet flap analysis while the Aerodynamics team designed the flap system configuration, which is shown in Figure 5.4.1.
Figure 5.4.1 - Cross Sectional View of the Blown Flap System.
As can be seen in Figure 5.4.1, there are three main components that were configured together inside the wing: the DP engine, the inlet to the engine, and the flap. Due to the flattened upper surface of the supercritical airfoil, the engine had to be placed high in the wing for the exhaust to blow over the top of the flap without ducting. This effect was amplified by the position of the rear spar. The flap itself was then shaped from the remaining portion of the airfoil. In an
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attempt to ensure attached airflow over the flap, a spoiler flap is actuated slightly downward while the blown flaps are in use.
5.5 Inlets Initially, the inlet was placed in the lower surface of the wing to allow for passive flow control. At high AOA’s, when the aircraft is trying to generate large amounts of lift, air would enter the inlet more directly and facilitate the distributed propulsion blowing system. However, an underside inlet is also at higher risk of foreign objects and debris (FOD) ingestion, especially on unprepared airfields such as a 328 ft soccer pitch. An advantage of moving the inlet to the upper surface of the wing is the straightening of the S-duct from the inlet to the engine. S-duct configurations deplete energy from the air, quickly defeating the passive flow control advantages from a below wing duct in this case. The inlet geometry was chosen to be an NACA type because of its efficient low speed performance. An example of an NACA inlet is presetned in Figure 5.5.1.
Figure 5.5.1 - Example of a NACA Inlet6
This is a very common flush mounted inlet originally designed for jet engine intakes, but now used for a multitude of inlet applications. The exit of the inlet was chosen to be the diameter of the DP engine and was designed with an aspect ratio of 3, as suggested by the NACA. A short length of ducting from the square inlet exit to the round engine was then needed. This length of ducting was the driving factor in the placement of the inlet on the wing surface. As suggested by the NACA,
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the inlet ramp was designed to have a slope of 7°, however it was made flat unlike the slightly curved design of the original6. Being flat, the ramp could be actuated up flush with the wing surface to aerodynamically seal the inlet in cruise.
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6. PROPULSION
6.1 Distributed Propulsion System The Distributive Propulsion concept is based upon using a series of small engines set across the wing. This is an alternative configuration very different from the conventional large under slung turbofan engines.
6.1.1 Principle
Distributed Propulsion spreads the engine exhaust across the wing to energize the boundary layer and increase the mass flow over the blown area. This effect can be achieved by mounting the engines internally or externally in close proximity to one another. An example of an externally mounted distributed propulsion design is presented in Figure 6.1.1.
Figure 6.1.1 - Distribute Propulsion Design3
The primary reason for selecting this unconventional configuration was the increase in lift offered relative to an un-blown wing. Research conducted by AVID3 explored the effects of using distributed propulsion in conjunction with Upper Surface Blowing (USB); the results of the study
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are found in Figure 6.1.2. A maximum wing of approximately 6 was achieved at a span-wise blowing distribution of . This is a substantial gain over an unblown airfoil.
Figure 6.1.2 - DP Impact on Lift Coefficient Distribution3
By increasing the total attainable lift, the aircraft is able to fly at significantly lower controlled speeds. Paired with an adequate primary propulsion system, the aircraft is more STOL capable. The benefits due to a DP configuration are clearly an attractive option for the C-328.
6.2 Jet Flap Theory and Integration In order to quantify the effect of the DP system, Jet Flap Theory (JFT) was used to predict the lift and pitching moment across the wing56.
6.2.1 Principle
The principle behind JFT is to expel a jet of air out of the trailing edge of the airfoil. The shear layer of air creates a pressure differential which acts as an extension of the total chord. This allows for increased lift from the jet flap or blown airfoil section and a rearward shift in the center of pressure.
Figure 6.2.1 illustrates the main parameters which were defined using the JFT. Namely the angle of attack ( ), flap deflection angle ( ), jet-momentum ( ), and the wing-chord-to-flap- chord ratio ( ).
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Figure 6.2.1 - Diagram of a Jet-Flapped Airfoil56
The theory derived in References45,56 was used to predict the jet flap characteristics. The jet momentum coefficient, Cµ, was utilized to define the thrust required per unit blown span.
6.1
The overall 3D wing coefficient of lift was calculated using,
6.2
This equation accounts for numerous correction factors which include: lift aspect ratio (F), airfoil thickness (t/c fuselage cut-out area ( ). Incorporating all of these presents an accurate), part spanestimation jet flaps of (λ)the andlift coefficient the which is tailoredν to the exact distributed propulsion and wing configuration chosen for the design.
6.2.2 Sizing
The type of jet flap used on the C-328 can be deflected to expel air out of the trailing edge as shown in Figure 6.2.1. This is different from traditional jet flaps, which cannot be deflected
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downwards. The benefit is greater lift potential from the wing, as the lift curve slope ( l/ is
45,56 steeper than that of a traditional jet flap as shown by Reference . ∂C ∂δ)
The chord-wise size of the jet flap was defined by choosing a equal to 0.3. This ratio was chosen based on structural limitations and maximum attainable flap sizes45,56. A larger ratio would result in ground interference at full flap deflection, while any smaller ratio would reduce lift potential from the wing.
The span-wise distribution of the jet flap was dictated by the wing geometry as shown in Figure 6.2.2. Both the fuselage and outboard aileron sections are not blown, as it was deemed control of the aircraft may be adversely affected. The rest of the wing was blown to maximize the wing’s lifting potential which resulted in 50.9% of the wingspan. The shaded area of Figure 6.2.2 illustrates the blown span of the wing.
Figure 6.2.2 - Blown Span/ Jet-Flapped Wing
The overall pitching moment of the aircraft changes when the DP system is operated. This occurs due to the change in lifting force across the blown area of the wing. Using Spence’s paper45, the induced pitching moment was found to shift 7.7% mean aerodynamic chord (MAC) rearward.
6.2.3 Stage Flap Deflection
For the C-328, JFT defined a required flap deflection angle of 62° at take-off in conjunction with a wing angle-of-incidence equal to 5°. Under this condition, the recovered horizontal thrust during takeoff was calculated. The net drag from the flap assembly is significant after a 30° flap deflection. Thus a staged flap deflection strategy was implemented to optimize the takeoff run and take advantage of the recovered thrust from the DP. The flaps progressively deflect until the
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required 62° deflection just prior to takeoff. The net recovered forward thrust is shown in Figure 6.2.3 for various flaps deflections.
Figure 6.2.3 - Net Forward Thrust from Distributed Propulsion at Various Flap Angles
6.3 Distributed Propulsion Design
6.3.1 Installation
The installation of the DP engines was dictated by the mission. The ESTOL requirement necessitated additional propulsion in the form of two large external turbofan engines (see Section 6.4). This meant that during cruise either the external turbofan engines or the distributed propulsion system would remain inoperable. Wind-milling drag for turbofans is a function of Mach number, bypass-ratio, and engine size. Since the main turbofans have to be fairly large, the penalty inquired by them was deemed impractical. Therefore, the decision was made to shut down the DP system during cruise and only use the main turbofans. The DP engines are internally installed as opposed to being on top the wing in order to reduce the drag and sfc penalty that would otherwise result.
6.3.2 Number of Engines
The number of DP engines dictates the magnitude of the thrust over the blown span. The JFT was then used to determine the thrust-per-unit span and the max attainable over the wing, ultimately resulting in a take-off velocity value.
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A case study of different engines was used to determine the most appropriate engine for the C-328. The study found that very small turbofan engines such as the Williams FJ22 and Pratt & Whitney PW610A would not be feasible for this design because the required thrust could not be met. The limitation of the number of engines that could fit within the wing dimensions led to the selection of larger sized turbofans with higher thrust outputs per engine. Prior to the final engine selection, further structural limitations were identified as illustrated in Figure 6.3.1.
Figure 6.3.1 - Engine Size Limitation
An over elongated engine would interfere with the wing’s mid-spar, while a large diameter engine would minimize the spar webbing causing further structural implications. The final engine selection would ultimately be limited by the number of engines that could be structurally supported by the wing.
6.3.3 Final Engine Selection
Several engines within the thrust class of 2,000-5,000 lbs were evaluated for the DP selection. This class was chosen since most of the engines were the right size to integrate into the wing as described in Section 6.3.2. A matrix was composed to evaluate how many engines could geometrically fit within the wing and how many engines were required for ESTOL according to the Jet Flap Theory (Section 6.2). In addition, the take-off lift coefficient and velocity were computed for each trial. This data was evaluated by the performance team to further evaluate ESTOL capabilities. The engine candidates and their performance results are presented in Table 6.3.1.
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Table 6.3.1 - Distributed Propulsion Decision Matrix
The final DP configuration selected consisted of 36 GE/Honda HF120 engines. This class of engine proved to be the best compromise between engine size and thrust output. The HF120 utilizes the most advanced technology in its engine class. It is designed for sustained performance with many enhanced durability features and a time between overhaul of 5,000+ hrs. GE/Honda state that there is no need for interim hot inspections and that the engine stays on wing 40% longer than competitors1. Additionally, the HF120 is an “off-the-shelf” engine making acquisition easier and maintenance costs cheaper. A picture of the HF120 is shown in Figure 6.3.2 and the accompanying engine characteristics are presented in Table 6.3.2.
Table 6.3.2 - HF120 Engine Characteristics1
390 Weight (lb) Fan Diameter (in) 16
Bypass Ratio 2.9
Pressure Ratio 24
Figure 6.3.2 - Engine Characteristics1 Cruise SFC (lb/lbf-hr) 0.6
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6.3.4 Performance
The HF120 DP system’s performance is summarized in Table 6.3.3. The summary shows performance at sea-level and hot & high conditions.
Table 6.3.3 - Summarized Distributed Propulsion Engine Performance Sea-Level Condition Hot & High Condition PERFORMANCE
Cμ 1.75 1.00
Engine No: 36 36
CLwing 4.21 3.47
VTO (ft/s) 86.3 106.3
Thrust at TO* (lbf) 59,040 40,150
SFC (lb/lbf-hr) 0.401 0.478
*Thrust at Take off supplied by performance prior to Short-Takeoff
The effective thrust curves of the C-328 engines were determined using ACS, GasTurb11, and compared to published results to verify the accuracy of the computations. The standard hot & high conditions (4,000 ft and 95 ) were estimated by using conditions at 10,000 ft Standard Temperature and Pressure (STP), which is approximately a 68% thrust correction factor from sea level. Hot & high conditions greatly degrade the thrust and sfc, which consequentially hurts the overall performance of the aircraft. In particular, the lift attained by the wing reduces in proportion to the reduction of the effective blown thrust.
The internal installation of the DP means some ducting is required. Due to ducting losses, a conservative ‘effective thrust’ coefficient was estimated at 0.8. However, during the detailed design phase a much cleaner ducting design was defined (Section 5.5). This means the ‘effective thrust’ coefficient is likely to increase, resulting in better DP performance. The improved performance has not been assessed as designed analysis of the ducting would be needed to fully ascertain the ducting losses of the final design.
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6.4 Cruise Engine System
6.4.1 Engine Selection
The primary objective of the main cruise engines was twofold. Firstly, since the DP would be shut off after take-off, the main engines had to provide an ample amount of power throughout all operating regimes including cruise at Mach 0.8. Secondly, they had to produce enough supplemental thrust to the DP engines to achieve the required take-off velocity for the ESTOL mission.
A high-bypass turbofan was chosen over a low-bypass or turbojet option as it is the most efficient for a long range transport with an operational region of Mach 0.8. A study was conducted to compare the major advantages and disadvantages of a two versus four engine design. A three engine configuration was considered but ruled out due to the large nose-down pitching moment it created and complexity of embedding the engine in the tail. The engine number study compared the Pratt & Whitney PW2000 and PW4090 based on their thrust size ratio 1:2. A summary of the key results are presented in Table 6.4.1 and an illustration comparing two and four engines is presented in Figure 6.4.1.
Table 6.4.1 - Two Versus Four Engine Design Comparison50 Two Engines Four Engines Weight (%) (PW4090) (PW2000)
Thrust per engine (lbf) N/A 90,000 45,000
Ground Clearance (ft) 15% 7.68 8.92
Total Engine Cost/Aircraft (Million $) 10% 24.0 33.2
Engine Maintenance Cost (Million 10% 72.5 79.6 $/year)
Static Thrust-to-Net Weight Ratio 40% 5.38 4.39
Total Cruise SFC (lb/lbf-hr) 25% 1.15 1.36
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Figure 6.4.1- Two versus Four Engine Design Illustration
The four engine configuration benefits from a lesser yaw moment that occurs in one engine inoperative (OEI) conditions, ground clearance for FOD consideration, and aircraft survivability due to multiple engine failure. However, a two engine configuration has a lower subsystem’s weight and ultimately a much higher aircraft static thrust-to-net weight ratio, which is a very favorable option for the ESTOL mission. Additional benefits include lower sfc, cheaper initial purchase cost, and cheaper operational costs. All advantages were weighed (Table 6.4.1) and the two engine configuration was chosen as best suited for the overall JFTL mission.
The thrust of the cruise engines was established by the thrust required during the ESTOL mission and at cruise. They were intentionally over-sized for sea level conditions to account for the thrust loss at hot and high conditions which has been identified as our main operational environment. The engines considered for the C-328 are shown in Table 6.4.2.
Table 6.4.2 - Candidate Cruise Engines50
RR Trent 895 GE-90-90B PW 4098
Static Thrust (lbf) 93,400 90,000 98,000
Static Thrust/Weight 7.13 5.41 6.06
Fan Diameter (in) 110 123 112
Bypass Ratio 5.79 8.4 5.8
Cruise SFC (lb/lbf-hr) 0.575 0.55 0.56
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Each candidate’s performance specification was weighed with emphasis primarily on total T/W, sfc, and the fan diameter which correlates to ground clearance. The Rolls-Royce Trent 895 was selected as it excelled in the majority of these categories. It is the industry leader for reliability and maintenance cost, while having the highest T/W in its class50. An illustration of the Trent 895 is presented in Figure 6.4.2 and the key characteristics are tabulated in Table 6.4.3.
Table 6.4.3 - Rolls Royce Trent 895 Engine Characteristics50
Static Thrust (lbf) 93,400
Weight (lb) 13,100
Fan Diameter (in) 110
Bypass Ratio 5.79
Pressure Ratio 40.7
Cruise SFC (lb/lbf-hr) 0.575 Figure 6.4.2 - Rolls Royce Trent 895 Engine50
6.4.2 Nacelle Design
The C-328 features a non-traditional nacelle design. The selection of a long-duct nacelle over the traditional separate flow nacelle was driven by the ESTOL requirements. Featured in Figure 6.4.3, the long duct nacelle has one exhaust nozzle, which forces the thermodynamic mixing of the fan and core gases. By equalizing the two gas streams, the engine gains a cycle performance advantage, which correlates to a higher propulsive efficiency, significantly lower sfc, jet noise, and infrared reduction, and greater reverse thrust capabilities. The disadvantages include added weight and higher installed drag. However, recent advances in engine technology, materials, and analytical methodology have nearly eliminated these handicaps.
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Figure 6.4.3- Long Duct Nacelle22
The dimensions of the nacelle were calculated using the methods found in Johnson22 and the results are presented in Figure 6.4.4 and Table 6.4.4. The inlet design parameters are based on the dimensions of the engine and representative ranges for transport aircraft gas turbofan installation. The nacelle design is tailored specifically to the Trent 895 engine in order to minimize thrust loss and spillage, as this would have an adverse effect on the critical phases of take-off and landing.
Table 6.4.4 - Nacelle Dimensions
Max Diameter (Dmax) 134.0
Exit Diameter (Dexit) 110.0
Cowling Diameter (DHL) 112.5
Internal Lip Diameter (DTh) 100.5
Diffuser Length (Ldiff) 172.0
Nacelle Weight (lb) 4,450
* All units in inches unless specified Figure 6.4.4 - Nacelle Dimension Nomenclature22
6.4.3 Reverse Thrust
In order to ensure a balanced field take-off and landing, the C-328 must stop in less than 328 ft, therefore reverse thrust capabilities are essential. As seen in the nacelle comparison chart in
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Figure 6.4.5, the long duct nacelle has approximately 30-35 % more reverse thrust capabilities than a typical design. At the point of touch down, the C-328 will have nearly 94,000 lb of reverse thrust at its disposal.
Figure 6.4.5 - Reverse Thrust Comparison22
6.4.4 Positioning
The horizontal position of the engines was based on experimental data and case studies. Ideally, the engine should be placed as far inward as possible to reduce OEI effects. However, closer placement suffers from the superposition of induced velocities from the fuselage and nacelle. This relation is demonstrated in Figure 6.4.6. The blue shaded region highlights the interference drag for conventional aircraft identifying the horizontal positioning for both a high-wing military C-17 and a low-wing civil Boeing 777. As can be seen, the C-328’s design conforms to best practice and its engines are located at 11.7 ft from the fuselage.
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Figure 6.4.6 - Horizontal Engine Placement23
Figure 6.4.7 illustrates the challenges faced with forward engine placement. Traditionally, the engine sits in front of the leading edge so that in the event of a severe engine malfunction it does not severely damage the wing. Foreign objects and debris was also highlighted as a threat which could potentially damage or destroy an engine. To prevent such occurrences, maximum ground clearance was a major design consideration. The Computational Fluid Dynamics (CFD) based design approach presented in Figure 6.4.7 reveals the relationship between the distance upstream of the wind and below the leading edge. As the proximity of the engine to the wing decreases, it must be placed further forward to avoid interference. Thus, a positioning compromise was found which places the C-328’s engines 17.7 ft forward of the leading edge with only a 2 ft gap under the wing. The final placement gives a total ground clearance of 7.68 ft. This is a very reasonable configuration as it is only marginally less than that of a C-17 at approx 8.9 ft and much greater than that of a Boeing 777 at 4.1 ft.
Figure 6.4.7 - Forward Engine Placement23
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6.4.5 Performance
The Trent 895 performance is summarized in Table 6.4.5.
Table 6.4.5- Summarized Distributed Propulsion Performance
PERFORMANCE Sea-Level Hot & High Cruise
Net Thrust (lbf) 95,000 64,067 10,277
SFC (lb/lbf-hr) 0.339 0.324 0.643
Figure 6.4.8 shows sfc versus thrust curves for different operational altitudes at the cruise condition of Mach 0.8. The red dot indicates the cruise point for the C-328 base upon optimized sfc at the required . Figure 6.4.9 shows sfc versus thrust curves for sea level and hot and high altitude combinations.
Figure 6.4.8 - Trent 895 SFC Thrust/Altitude Curves
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Figure 6.4.9 - Trent 895 SFC Thrust Curves for Hot and High and Sea Level Conditions
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7. STRUCTURES
This design required the integration of the new distributed propulsion technology, which drove many of the design decisions.
7.1 Velocity-Load Diagram The Velocity-Load (V-n) diagram shown in Figure 7.1.1 illustrates the structural limits of the C-328 at given speeds. Stall, cruise, dive speeds and the flap-down maneuver envelope are all marked on the plot. The maneuver limits of the Ostrich, 3.0 and -1.5, are typical of a military transport35. The critical load factor of 3.2 occurs at a wind gust of 50 fps during cruise. The 50 fps gust line can be seen in blue. Red dashed lines represent the other gust conditions.
These values were calculated using the methods described in Johnson21. The aerodynamic stall curves were calculated using the following equations21:
(for stall) 7.1
(for inverse stall) 7.2
and were obtained by projecting the maximum and minimum lift and drag coefficients onto the axis parallel to the weight of the aircraft (vertical axis), respectively. The gust load factors where determined using21:
7.3 where V is the velocity of the aircraft and U is the velocity of the wind gust.
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3.5 nmax = 3.2 3 Flaps Down 2.5 50 fps 25 fps 2 66 fps 1.5 Vdive = 693 knots 1 V = 462 knots 0.5 cruise Vstall = 222 knots Load Factor, n Factor, Load 0 0 100 200 300 400 500 600 700 800 -0.5
-1 V = 492 knots -1.5 neg stall Equivalent Airspeed, knots
Figure 7.1.1 – V-n Diagram
7.2 Wing Box Layout The greatest structural challenge in designing the C-328 was the wing. Early in the design process, it was decided to place the DP engines inside the wing to eliminate the drag penalty at cruise, and it quickly became apparent that this would drive the design of the wing box. Figure 7.2.1 shows the integration of the wing box with the wing. The wing is positioned on top of the fuselage, allowing for a continuous structure.
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Figure 7.2.1- Wing Box Structure showing Distributed Propulsion Engines
7.2.1 Spars and Ribs
The front spar is positioned at 15% of the chord for the entire length of the wing box, which extends from root to the tip of the aileron. The rear spar is positioned at 70% of the chord from root to the tip of the DP compartment, and is positioned slightly aft of 70% of the chord for the remainder of the wing box. The mid spar is positioned at 37.5% of the chord for the entire length of the wing box.
The position of the wing box was driven by the placement of the DP engines and the sizing of the flaps. The DP engines were originally positioned in front of the rear spar, moving the rear spar to approximately 75% of the chord; however, the added risk of structural failure due to the heating of the rear spar from the exhaust of the DP engines drove the placement of the DP engines aft of the rear spar. A diagram of the wing box layout can be seen in Figure 7.2.2.
Web thicknesses at the root and tip were determined for each spar using methods found in Loughlan26,27 and Niu35. The resulting values for this aircraft can be found in Table 7.2.1.
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Table 7.2.1 - Web Thicknesses at Root and Tip for Wing Box Spars
FRONT CENTER REAR
in in in
ROOT 0.853 1.070 0.902
TIP 0.116 0.116 0.109
Rib spacing was driven by the diameter of the DP engines. A distance of 30.1 in allows space for the engine intake holes in the rear spar with an additional 1 in on either side. This spacing results in 29 ribs in each wing.
FRONT SPAR Aileron MID SPAR REAR SPAR
FLAP STRUCTURE
DP Engines
Figure 7.2.2 – Wing Box Layout
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7.2.2 Integration of Distributed Propulsion
The DP engines are connected to the rear spar using A-frames, which are aligned with the ribs. Air is directed to the engine intakes via ducting from the upper surface through the rear spar. More information on the ducting can be found in Section 5.5.
The resulting configuration at the root can be seen in a cross-sectional diagram in Figure 7.2.3. The engine is mounted in the top half of the airfoil to maximize the blowing effect, minimize ducting from the upper surface, and allow space below for the flap activation mechanism. A spoiler on the upper section of the airfoil is deflected up during the short landing to act as an airbrake, and deflected down during the short takeoff to direct the exhaust from the DP engines over the flap.
Ducting DP Engine
Flap Retracted
Front Spar Mid Spar Rear Spar
Flap Deployed
Figure 7.2.3 - Cross-Sectional Diagram of Wing at Root
The configuration at the tip of the DP engine section can be seen in Figure 7.2.4. The engine at this cross-section fills the area between the upper and lower skins of the airfoil. It should also be noted that the ducting and intake for the DP engine at this cross-section extends from the mid spar to the rear spar.
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Ducting DP Engine
Flap Retracted
Front Spar Mid Spar Rear Spar
Flap Deployed
Figure 7.2.4 – Cross-Sectional Diagram of Wing at Tip of Distributed Propulsion Section
7.2.3 Flap Attachment
The flaps are attached to the same A-frames at a hinge joint. A hydraulic actuator is used to deploy the flaps. At each flap attachment point, the A-frame extends below the lower skin of the wing into a fairing. A rough diagram of this mechanism can be seen in Figure 7.2.3 and Figure 7.2.4.
7.2.4 Finite Element Model
A finite element model (FEM) of the wing box was created in ANSYS. The meshed model can be seen in Figure 7.2.5. The FEM consists of 4658 nodes, 5466 elements, and 27888 degrees of freedom. The model includes the spars, ribs, upper and lower skins, and the A-frames which support the DP engines and flap structure. The root displacement is constrained to zero in all degrees of freedom in order to model the wing box as a cantilever beam. This allows the wing the forces acting on the wing to be considered independently from the rest of the aircraft. The engine weight is transferred to the wing box through two inelastic elements. This allows for accurate engine load placement, while still considering the effect of the load on only the wing box, eliminating the complexity of the engine pylon.
The lift distribution was provided by the Aerodynamics group, and can be found in Figure 5.2.3. The majority of the lift in cruise occurs behind the rear spar on the flap structure, and forward of the front spar. The FEM models the lift with point loads applied to each A-frame behind the rear spar and point loads applied to inelastic elements connected to the front spar.
Initial model tests revealed that the lift distribution requires large A-frames to transfer the load from the flap structure to the wing box. When the weight of the fuel was added, the model analysis would not execute; therefore, the model cannot be verified. In the future, the maneuver
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load cases will be tested, and the FEM will be used to determine which wing box locations will need additional stiffening.
Figure 7.2.5 – Meshed Finite Element Model of Wing Box
7.3 Other Structural Components
7.3.1 Fuselage
The frame spacing in the fuselage was derived from the following equation35:
7.4 where E is the modulus of elasticity of Aluminum 6061-T6, I is the moment of inertia of the frame, D is the diameter of the fuselage, M is the bending moment on the fuselage (provided by the Aerodynamics group), and L is the frame spacing. An iterative method was used to find the optimum frame spacing. The spacing is 22.0 in, resulting in a total of 69 frames. A diagram of the frame spacing can be seen in Figure 7.3.1. The pressure bulkhead locations are also noted in the figure. The bulkheads are located forward of the cockpit and aft of the cargo door. A third, smaller
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pressure bulkhead is located behind the cockpit to seal it from atmospheric conditions. The cargo hold is unpressurized.
Figure 7.3.1 – Top Down View of Fuselage showing Frame Locations. Arrows indicate Pressure Bulkhead Locations
7.3.2 Horizontal and Vertical Stabilizers
The horizontal and vertical stabilizer structures consist of a front spar at 15% of the chord, a rear spar at 75% of the chord, and a rib spacing of 24 in. The horizontal stabilizer torque box includes an additional mid spar at 45% of the chord to assist in carrying the added load due to the large size of the stabilizer. The vertical stabilizer is attached to the fuselage at the rear pressure bulkhead.
7.4 Materials The majority of the aircraft structure is composed of conventional materials: spars and ribs are Aluminum 6061-T6 and 7075-T73; wing skins are Aluminum 2024-T3. The flap structure, which is exposed to the hot gas exhaust of the DP engines, must be constructed from a Titanium 140A or Titanium 155A alloy. The Boeing YC-14 uses a titanium alloy for a blown flap, and was able to sustain temperatures in excess of 800 °F while maintaining a very low material density, 0.174 lbs/in3. This low density is important in preventing additional stresses and moments from being placed on the surrounding structure34. The stiffeners used around the holes in the rear spar are also composed of a titanium alloy. The additional cost considerations of titanium alloys are discussed in Section 11. Graphite and epoxy composites compose portions of the vertical and horizontal stabilizer skins. While composites are not new, they are not widely used on military aircraft, excluding the vertical and horizontal tails. Due to the C-328’s large wing and DP technology, it was decided that the wing skin should be composed of the more traditional aluminum alloy sheets. Steel will be used in the construction of the large turbofan nozzles and the landing
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gear struts. Figure 7.4.1 shows a rough material weights breakdown based on the densities of the materials and the aircraft structure sizing. The “Other” materials category consists of plastics, Kevlar, and other minor materials.
Steel Other 7% 5%
Graphite/Epoxy 20% Aluminum Alloys 50%
Titanium Alloy 18%
Figure 7.4.1 – Material Weight Breakdown
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8. WEIGHT AND BALANCE
8.1 Empirical and Group Methods Used During the design process, two methods were used to calculate the aircraft’s weight and CG. In the early stages of design, rough estimates were made based on the percentage of gross weight, as well as planform and wetted areas. As the design progressed and the sizing and layout of the aircraft evolved, a more detailed statistical group weights method was implemented. The latter method utilized regression analysis to form statistical equations since the exact weight for each component was unknown39.
During the initial phase of the design, each group needed weight approximations. After calculating the planform areas of the wing and empennage, empirical data39 was referenced to approximate the respective weights for each item in Table 8.1.1. The exposed planform areas were multiplied by the referenced empirical data to arrive at weight estimations of major structural members.
Table 8.1.1 - Empirical Data for Transport Aircraft for Approximate Empty Weight Buildup39
Item lb/ft2 Multiplier
Wing 10.0 Sexposed planform
Horizontal Tail 5.5 Sexposed planform
Vertical Tail 5.5 Sexposed planform
Fuselage 5.0 Swetted area
Landing Gear .043 TOGW
The wetted area of the fuselage was calculated using geometric approximations of the fuselage shapes. The nose, main section and tail areas were calculated using cylindrical areas24 and scaling multipliers suggested by Raymer39.
Once estimations for the C-328’s operating weight empty (OWE) were completed, a detailed analysis was performed to more accurately calculate the OWE and CG. A statistical group weights
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method was implemented to achieve this. In lieu of custom regression analysis, equations provided in Raymer39 were utilized. The results of this analysis are shown in Table 8.1.2.
Table 8.1.2 – Component Group Weights and Moments
Weight x-bar x-Moment Groups Pounds Feet Foot-Pounds W x W*x Wing 22210 41 907344 Horizontal Stab 6802 161 1094268 Vertical Stab 4092 150 612120 Fuselage 18344 47 870034 Main L.G. 15505 57 878312 Nose L.G. 2584 9 23139 Main Nacelles 8895 33 291471 Nacelle Group--DP 3443 64 220584 Structure Weight 81875 561 4897272 Engine Controls 23 35 802 Starter (pneumatic) 289 35 10022 Fuel System 2731 35 94741 Flight Controls 898 35 31155 Instruments 237 16 3883 Hydraulics 363 43 15707 Electrical 2742 43 118703 Avionics 1693 43 73296 Furnishings 2301 43 99626 Air conditioning 1910 43 82684 Anti-ice 661 40 26759 Handling Gear 99 54 5359 Cargo Handling 1860 54 100480 Engine Controls 167 76 12670 Starter (pneumatic) 234 76 17697 Fixed Equipment 16207 672 693585 Main Engines 28359 27 756686 Distributed Engines 21579 64 1382656 Propulsion 49937 91 2139342 Empty Weight 148020 1324 7730199 Crew 441 10 4266 Stryker Payload 66000 52 3403066 Zero-Fuel Weight 214461 1385 11137531 Stryker Mission Special Forces Payload 26400 52 1361226 Zero-Fuel Weight Special Forces Mission 174861 1385 9095692
Fuel--Stryker Mission 81336 52 423743 TOGW--Stryker Mission 295797 1437 11561274
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Fuel--Special Forces Mission 131721 52 6862350 TOGW--Special Forces Mission 306582 1437 15958042 Stryker Mission x_cg (ft) Full Fuel and Payload 51.98 Full Fuel and No Payload 52.10 No Fuel and No Payload 52.10 Special Forces Mission x_cg (ft) Full Fuel and Payload 52.05 Full Fuel and No Payload 52.10 No Fuel and No Payload 52.10
From this table it can be seen that the C-328 has an OWE of 148,020 lbs, with more than 55% of that weight coming purely from the aircraft structure itself. It also may be seen that the TOGW for the Special Forces Mission is approximately 11,000 lbs heavier than for the Stryker Mission. The reason the TOGW is greater for the mission with the smaller payload is the fuel required to carry out the mission to completion (see Section 11.3). As shown above, more than 50,000 lbs of extra fuel are required for the Special Forces Mission as compared to the Stryker Mission, since the operating radius for the Special Forces Mission is so much larger.
8.2 Center of Gravity The CG of the aircraft was calculated by summing the moments of each aircraft component about the nose of the aircraft. While stationary on the ground, the resultant CG location was 52.10 ft aft of the nose of the aircraft. Figure 8.2.1 shows this location on the aircraft.
Figure 8.2.1 – Side View with CG Location
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When the C-328 performs the AJACS and Special Forces missions, weight is lost due to fuel burn. This required the determination of the CG envelope. The CG envelopes for both missions are shown in Figure 8.2.2 and Figure 8.2.3. A key result presented by these plots is the small variation in CG location through the range of possible aircraft weights. The CG envelope was optimized by situating the fuel tanks and payload directly over the CG of the empty aircraft. This allowed for payload and fuel CG neutrality.
AJACSStryke r Mission
310000
290000
270000
250000 Varying Fuel, Full Payload 230000 Varying Payload, Full Fuel
210000 Varying Fuel, No Payload
Weight (lbs) Weight Varying Payload, No Fuel 190000
170000
150000
130000 51.92 51.94 51.96 51.98 52.00 52.02 52.04 52.06 52.08 52.10 52.12 C.G. (ft)
Figure 8.2.2 - Potato Plot showing the CG Envelope of the AJACS Mission.
Special Forces Mission
350000
300000
250000 Varying Fuel,Full Payload Varying Payload, Full Fuel Varying Fuel, No Payload 200000 Weight (lbs) Weight Varying Payload, No Fuel
150000
100000 52.01 52.02 52.03 52.04 52.05 52.06 52.07 52.08 52.09 52.10 52.11 C.G. (ft)
Figure 8.2.3 Potato Plot showing the CG envelope of the Special Forces Mission
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9. STABILITY AND CONTROL
9.1 Tail Sizing The aircraft employs a T-tail configuration which provides two main advantages over other common tail types. First, the large moment arm for the elevator increases the effectiveness of the control surface in generating a nose up pitch moment. Second, the T-tail configuration prevents the horizontal tail from being blanketed by downwash from the DP flap system. Both the horizontal and vertical tail surfaces were initially sized using Raymer’s method of volume coefficients39.
9.1
9.2
The volume coefficients, and , were initially assigned values of 1.00 and 0.09, respectively. The vertical tail volume coefficient was changed to 0.0855 due to the T-tail end-plate effect which allows a 5% reduction. The horizontal coefficient was reduced to 0.95 since it is out of the downwash region, which also allows a 5% reduction in volume coefficient.
During the preliminary design phase, the horizontal tail was over half the area of the wing as sized. This was much larger than the horizontal tail to wing area ratio of similar medium transport aircraft. In an attempt to shrink the horizontal tail area, the fuselage was lengthened to create a longer moment arm.
The vertical tail sizing was also initially calculated using volume coefficients. Once the aircraft’s engines were selected, an OEI analysis was performed to check Federal Aviation Regulation (FAR) compliance. The OEI tail sizing was calculated using the parameters for the AJACS mission. Due to the low takeoff speed of 25.3 m/s in the ESTOL mission, the tail could not be sized for OEI without over stabilizing the aircraft during all other mission segments. Additionally, all main thrust is required in order to takeoff and land in the ESTOL mission. The resulting horizontal and vertical tail areas are specified in Table 9.1.1 below.
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Table 9.1.1 - Summary of Tail Surface Sizing
Surface Area (ft2) Span (ft) Chord (ft)
Horizontal Tail 1938 95.4 20.0
Vertical Tail 928 33.5 29.2
9.2 Control Surface Sizing The aircraft ailerons, elevator, and rudder were initially sized using volume coefficients39. The elevator area estimate from the volume coefficient method was considered acceptable when compared with existing transport aircraft; however, due to the aircraft’s large wing area, the ailerons were oversized. The area of the ailerons was reduced based on the available control power derived from deflection angles and moment arms. The rudder area was calculated based on control power with one engine out. The final control surface sizing is presented in Table 9.2.1 below.
Table 9.2.1 – Summary of Control Surface Sizing
Surface Area (ft2) Span (ft) Chord (ft)
Ailerons 548 22.6 7.9
Elevator 708 87.9 8.2
Rudder 278 24.3 11.5
9.3 Longitudinal Stability
9.3.1 Trimming for Takeoff
The pitching moment of the aircraft proved difficult to resolve during the detailed design phase. Initially, the configuration was stable; the AC of the wing was behind the CG. As a result, the lift of the aircraft produced a nose down pitching moment. The aircraft also experienced a large nose down aerodynamic pitching moment augmented by the distributed propulsion system
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and blown flaps. Combined, these two factors produced too much downward pitching moment to effectively trim with the horizontal tail. To address this problem, the wing was repositioned forward in order to place the AC of the wing ahead of the aircraft CG. This resulted in an unstable aircraft with a static margin of 12.5% MAC. The horizontal tail was also designed to have variable incidence for trim. The tail is actuated at -8° during takeoff, with an elevator deflection of 5° for trim. In cruise, the tail incidence angle is approximately zero.
9.3.2 Dynamic Longitudinal Stability
Stability and control derivatives were calculated, along with the aircraft’s mass moments of inertia, and used to assess the C-328’s handling qualities in different modes of flight. All CL values and wing aerodynamic moment data were calculated using Tornado40. The moments of inertia were calculated using an estimation method in Raymer39. The moments of inertia are shown in Table 9.3.1.
Table 9.3.1 – Mass Moments of Inertia
Mxx (slugs/ft2) Myy (slugs/ft2) Mzz (slugs/ft2)
5,055,386 12,991,592 14,406,930
The resulting parameters from these analyses, such as damping ratios or natural frequencies, were compared against requirements set forth by the military42. The military airworthiness regulations for airplane performance define three different levels of flying quality for different classes of aircraft and different phases of flight. The Ostrich qualifies as a Class II transport for the Special Forces mission and as a Class III transport for the AJACS and JHL missions. Stability modes were evaluated for Phase B, cruise and Phase C, takeoff and landing. All tables presenting a comparison of military requirements and Ostrich data are comparisons to level one flying requirements. Level one flying quality is the highest level attainable demanding little to no compensation from the pilot for the aircraft’s handling. Level two flying quality has decent handling but requires some compensation by the pilot. Level three flying quality demonstrates poor handling and requires extreme compensation by the pilot. For longitudinal dynamic stability, two modes were considered – the short period and the phugoid. The short period damping ratio limits, as specified by military requirements42, are
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presented in Table 9.3.2 and the phugoid minimum damping ratio requirements are presented in Table 9.3.3. The short period analysis resulted in a high natural frequency due to large wing area. Comparison to the military requirements show the aircraft has poor flying qualities in the short period mode. The aircraft also has poor handling qualities for the phugoid mode due to the large horizontal tail. It experiences level 2 flying quality for takeoff and level 3 flying quality for cruise. Level 3 flying qualities are unacceptable. To compensate for the aircraft’s substandard flying quality, a stability augmentation system will be used.
Table 9.3.2 - Short Period Damping Ratio Limits.
Phase C Phase B
Level I MIL 0.35 – 1.30 0.30 – 2.00 Requirement
Calculated 3.43 0.21
Table 9.3.3 - Phugoid Minimum Damping Ratio
Phase C Phase B
Level I MIL 0.04 0.04 Requirement
Calculated 0.29 0.035
9.4 Lateral-Directional Stability Most lateral directional stability and control derivatives were calculated using LDstab.exe, a program written by Joel Grasmeyer and methods prescribed in Roskam41. These derivates are shown in Table 9.4.1.
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Table 9.4.1 - Lateral Directional Stability and Control Derivatives
Takeoff Cruise
Cyβ -0.76 -0.79
Clβ -0.36 -0.36
Cnβ 1.40 1.41
Cyda 0.0 0.0
Clda 0.046 0.046
Cnda 0.0064 0.0064
Cldr 0.016 0.017
Cndr 0.085 0.91
Cyr 0.78 0.78
Cnr -0.55 -0.55
Clr 0.12 0.12
Three dynamic modes were analyzed, the spiral mode, dutch-roll mode, and roll mode. The spiral mode analysis is summarized in Table 9.4.2. According to this calculation, the spiral mode of the aircraft is stable. It produces level 1 flying qualities for cruise, but takeoff flying quality suffers due to a very small time to double amplitude. This is caused by large stability derivatives from the change in rolling and yawing moments resulting from a disturbance in sideslip angle, . These derivatives, and are large because of the dihedral effect, the large area of the wing, and the very high created by USB.
Table 9.4.2 - Spiral Mode Minimum Allowable Time to Double Amplitude,
Phase C Phase B
Level I MIL 12 sec 20 sec Requirement
Calculated 11.42 sec 0.28 sec
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The dutch roll analysis is summarized in Table 9.4.3, Table 9.4.4 and Table 9.4.5. This analysis shows that the dutch roll mode is also stable. The military requirements42 specify criteria for undamped natural frequency and damping ratio, as well as their product, referred to as the real root part value. The aircraft experiences level 1 flying qualities for cruise and level 2 flying qualities for takeoff. During takeoff, the damping ratio is too small and does not meet level 1 criteria. The small damping ratio is a result of a large undamped natural frequency. Both real root part values for takeoff and cruise satisfy level 1 criteria.
Table 9.4.3 - Dutch Roll Undamped Natural Frequency, ωnD
Phase C Phase B
Level I MIL 0.4 sec 0.4 sec Requirement
Calculated 2.66 sec 17.66 sec
Table 9.4.4 - Dutch Roll Damping Ratio,
Phase C Phase B
Level I MIL 0.08 0.08 Requirement
Calculated 0.06 0.40
Table 9.4.5 - Dutch Roll Real Root Part Value,
Phase C Phase B
Level I MIL 0.10 rad/sec 0.15 rad/sec Requirement
Calculated 0.16 rad/sec 7.11 rad/sec
The roll mode analysis is presented in Table 9.4.6. The aircraft has a stable roll mode and meets level 1 flying quality requirements in both takeoff and cruise. Roll rate coupling stability was also checked and found to be unstable. However, the maximum roll rate attainable by the
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aircraft is 13.75° per second which is less than the minimum roll rate needed to induce instability.
Table 9.4.6 - Roll Mode Maximum Allowable Tim Constant,
Phase C Phase B
Level I MIL 1.4 sec 1.4 sec Required
Calculated 1.03 sec 0.02 sec
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10. SYSTEMS
The function of the landing gear on this aircraft is to facilitate the repeated operation on unprepared airstrips for short take-off, landings and ground maneuvers. This type of operation puts a high demand on the landing gear as its performance is critical to the success of any required mission. Bearing this in mind, the following design is proposed for the C-328.
10.1 Landing Gear The landing gear for the C-328 was modeled after the gear from the A400M. The A400M landing gear is shown in Figure 10.1.1. This type of main gear system allows for a vertical retraction of the wheels, allowing for a much smaller space required to house the landing gear while they are not in operation. This advantage, coupled with the well-optimized strength to weight of this design, accounts for the decision to employ a similar main landing gear system.
Figure 10.1.1 - A400M Landing Gear57
10.1.1 Layout and Arrangement
The landing gear adopts a conventional tricycle configuration, consisting of a single nose gear unit along with two sets of main landing gear. This configuration allows for the optimal conditions for loading and unloading the cargo bay as well as a high AOA short-field landing. The main gear is arranged in what is commonly known as the tri-twin tandem arrangement, as shown in Figure 10.1.2.
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(Dimensions shown in ft)
Figure 10.1.2 - Tri-twin Tandem Landing Gear Arrangement
The reason for applying the tri-twin tandem arrangement for the main gear is mainly to achieve the required flotation for operations on soft fields by increasing the contact area. An alternative considered was track type gear; however, this idea was rejected because it introduces a weight penalty and complicates shock absorber design. The tri-twin tandem arrangement was chosen as the best solution in order to keep weight to a minimum while operating effectively on a soft surface such as wet grass.
The following steps were taken to determine the longitudinal location of the main landing gear units. The aft towing angle was set to 15°, as shown in Figure 10.1.3 below, in order to prevent the aircraft from tipping over when brakes are applied to produce a deceleration of 8 ft/s² or greater. For this particular aircraft, the aft towing is an important design consideration as the landing length is very short, resulting in the need for very high deceleration rates.
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Figure 10.1.3 – Aft towing angle
The angle between the rear-most gear and the aft fuselage becomes an important geometric constraint when considering the rotation action of the aircraft on take-off. To ensure that the CG does not rotate past the contact point of the rear-most gear, the main gear have been positioned so that the tail tipping angle, shown in Figure 10.1.4, is 16°. This avoids potential stall on takeoff, which can occur at angles greater than 20°.
Figure 10.1.4 – Tail Tipping Angle
The final stage in positioning the landing gear is to take the static gear loads into consideration. By placing the nose gear as far forward as possible this has the effect of maximizing flotation and stability.
10.1.2 Tires
Regarding tire selection, pneumatic tires have been selected for the following reasons. These tires are suitable for taxiing over rough surfaces and provide good adhesion with the runway surface, which is required and desirable for heavy braking and ground maneuvers. Low rolling drag of the tires is also vital in order to have short takeoff ability. However, to provide necessary friction, the tire tread will be ribbed, similarly to many aircraft capable of tactical
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landings, such as the C-130 and C-17. Due to relatively high TOGW, good flotation characteristics are desirable for efficient operation. Using the approach presented in Currey15, the loading on each tire has been calculated as shown in Figure 10.1.5 below.
(Loadings in lb)
Figure 10.1.5 – Individual Tire Loading
In order for the tires to handle the calculated loads, the tire pressures will be clearly specified because the loading is proportional to tire pressure. To adequately support the required loads, the tires will use a pressure of 100 psi in the main gear and 170 psi in the nose gear. Using high pressure tires offers reduced weight, rotational inertia, and cost. One disadvantage to extremely high pressure tires is the increased wear rate of the tire. A compromise has been made because the aircraft is not expected to perform a high number of landings in a short time frame. For conventional operation in which austere short-field landing capability is not needed, more conventional, lower-pressure tires can be used in order to increase the tire lifespan.
10.1.3 Main Gear Housing and Structure
The main landing gear is housed in two lower fuselage sponsons, shown in Figure 10.1.6. Despite the disadvantage of increased aerodynamic drag over “inside fuselage” main gear housing, the several notable convenience and weight advantages of the system outweigh the calculated drag penalty.
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Figure 10.1.6 - Sponson Configuration (front view)
The sponsons’ increased lateral ground spacing greatly increase the margin of stability for roll while the aircraft is still on the ground. Additionally, the sponsons are advantageous to the cargo loading and unloading scheme. Without sponsons, the ability to “kneel” the main gear is not feasible. Kneeling gear, which will be explained in further detail below, allow for easier cargo handling. External landing gear housing also allows more of the fuselage internal volume to be used for payload area. Finally, having external bays for landing gear allows for much larger and stronger landing gear systems without requiring significant increase in system weight.
Landing gear are an essential system to a successful mission, and therefore are in need of protection from any debris or projectiles which could cause damage. Sponsons in the design allow for protection of sensitive systems from mud and other ground debris on unprepared fields, as well as armored protection from hostile attack.
In order to properly design the landing gear, initial calculations of strut length were made based on equations presented in Currey15 and Conway13. These calculations indicate that the strut length for the landing gear should be at least 25 inches in all cases.
10.1.4 Special Features
The several special features of the landing gear system are essential in the successful operation of the Ostrich’s unconventional missions. These features and their advantages are outlined below. It is important to note, however, that these features add to the complexity and cost of the aircraft, as well as the weight of the landing gear in general. These penalties were
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taken into account when determining whether the benefits in fact outweighed the weight and cost/complexity penalties.
10.1.4.1 Kneeling To allow quick and efficient cargo loading/unloading capability, the landing gear will have the ability to ‘kneel’ thus reducing the angle of the cargo ramp door with respect to the ground. The aircraft has been designed with the ability to kneel 1.31 ft, which yields an angle between the cargo door and the ground of 8.8°. This is adequately shallow for efficient loading/unloading procedures.
10.1.4.2 Nose Gear Actuation To assist with takeoff rotation on a short field, the nose gear will also have the ability to extend further. This actuation will be commanded at a given takeoff run velocity to give additional incidence to the wing for short-field takeoff. Using nose gear actuation can increase the rotation angle of the aircraft by 2.8°. Based on aerodynamics calculations, this angle adds a performance benefit sufficient to justify the inclusion of this design feature.
10.1.4.3 Tail Bumper Based on the guidance given in MIL-L-87139, military aircraft are recommended to make use of a tail bumper. The tail bumper exists to prevent an AOA required for 90% of on the wing. For this aircraft, the angle of rotation required for the tail bumper to touch the ground (assuming main gear in static loading position) is 15°.
10.1.5 Pilot Control and Operation
Since a retractable gear has been adopted, indication for when the landing gear is up and locked or down and locked will be given on the flight deck. The landing sequence is fully automated and controlled via a simple lever positioned next to indication lights. The pilot will also be able to control the kneeling of the aircraft; however, the nose gear actuation on takeoff will be handled autonomously through the flight control system. In the case that the landing gear sequence does not complete, an aural warning will be given on the flight deck in compliance with MIL-S-9320 and MIL-L-87139. Steering of the aircraft for ground maneuvers is controlled via a steering tiller or through use of the rudder pedals, which controls the nose wheel direction. The control system for the steering of the aircraft complies with MIL-S-8812. The nose gear has the ability to turn up to 45° in either direction. In order to prevent scrubbing of the main gear tires,
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the struts on which they are mounted will have the ability to rotate up to 10°. This is a necessary feature for the aircraft to enable operation on soft turf fields.
10.2 Fuel Systems The fuel system of the aircraft includes six tanks, all of which are located in the main wing. Two of these tanks serve as reserve tanks for emergency situations. An additional two tanks exist for wingtip fuel dumping when necessary. Table 10.2.1 below gives specifications for the fuel system capacity.
Table 10.2.1 - Fuel Tank Capacity Specifications
Tank Depth (ft) Planform Area (ft²) Volume (ft³) Weight (lbs)
1 & 4 2.81 4.96 1,412 700,352
2 & 5 3.28 516 1,696 875,136
3 & 6 1.31 150 212 31,800
Total 1,162 3,320 1,607,288
The fuel tanks used in the aircraft are self-sealing bag fuel tanks. The tanks exist within the structure of the wing, allowing fuel to flow through holes in the rib structure. All analysis of the wingbox structure, therefore, had to be performed taking these fuel tanks into account. These tanks are generally considered to be the ideal tank type on an aircraft which may be subject to attack by small-arms gunfire. Because these tanks take up nearly a quarter of the volume of fuel they contain, the fuel tank dimensions listed above were sized such that there would be sufficient available room for the tanks and the fuel needed within the wing volume.
In order to accommodate the fuel needs of all the engines on board (both DP engines as well as the main turbofans), multiple fuel pumps exist. Additionally, transfer pumps would be necessary to allow for a digital fuel level management system. Included in the original requirements for the aircraft was in-air refueling capability. A probe (for Navy and European application) could be optionally installed toward the nose of the aircraft, or a port (for USAF refueling missions) through which fuel could be pumped through a system of fuel pipes to the main wing fuel tanks.
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11. PERFORMANCE
The C-328 is required to perform two highly differing missions. The AJACS mission requires the aircraft to carry a heavier payload for a shorter distance with a take off and landing distance of 1500 ft. The more challenging Special Forces mission, with a take off and landing distance of 328 ft, drove the performance requirements to a level not achieved by any of today’s aircraft. This required the use of distributed propulsion, a new technology and consequently a new and bespoke approach to modeling the performance of such an aircraft.
11.1 Powered Lift Mission Segments The powered lift mission segments are the segments in which the distributed propulsion was utilized. Using this new technology required performance analysis beyond the scope of most existing and conventional performance texts. The complex interconnections between the DP engines horizontal thrust, , and stall speed was established from JFT (section 5.2) and demanded the development of a custom performance analysis scheme. Therefore, custom short takeoff and short landing performance analysis codes were written, utilizing time-step integration to calculate the instantaneous acceleration of the aircraft at every step of the ground roll from the initial rest state to the takeoff velocity.
Figure 10.1.1 presents a free body diagram of the aircraft during the short takeoff and landing mission segments in both the AJACS and Special Forces missions. L
Tmain
TDP D Ffriction
W
Figure 11.1.1 - Free Body Diagram During ESTOL Ground Roll.
In the next two sections, the analysis for the Special Forces mission is presented in detail as this was the driving mission in the design of the takeoff and landing routines. At the end of
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each section, the differences in the takeoff and landing schedules for the AJACS mission are presented for comparison.
11.1.1 Short Takeoff Ground Roll
The Free body diagram shown in Figure 10.1.1 was used to create the expression for the instantaneous acceleration during the short takeoff ground roll presented in Equation 11.1.
11.1
The instantaneous velocity used in Equation 10.1 was calculated during each time step as a function of the acceleration. Additionally, , and distributed propulsion horizontal thrust, were calculated as functions of the flap deflection angle according to jet flap theory as presented in Section 5.2. The C-328 flap design allowed for the full thrust of distributed propulsion engines to be utilized at flap deflections of 30° or greater. At flap deflections less that 30°, the flap mechanism interferes with the exhaust of the DP engines. Additionally, a takeoff flap deflection angle of 62° was required to allow for maximum possible CL according to JFT.
To achieve optimum takeoff performance for the Special Forces mission, a variable flap deflection schedule was utilized. Varying the flap angle during the ground roll allowed for maximum horizontal thrust (and accordingly maximum acceleration) at the start of the ground roll as well as the required 62° flap deflection at the end of the ground roll. Without variable flap deflection, the C-328 was required to run the entire ground roll at 62° flap deflection. The large size of the flaps and the diminished horizontal thrust in this configuration made the C-328 unable to achieve the required horizontal acceleration during the early stages of the ground roll to meet the ESTO requirement. Although the variable flap deflection schedule adds a considerable degree of complexity to the ground roll, it was still incorporated as ESTO operation requires its performance benefits. Additionally, ESTO operation requires the C-328 to operate flawlessly to achieve the desired ground roll and accordingly, failures such as engine out and flap deflection failure would be catastrophic during ESTO operation. During further stages in the design process, the reliability of these systems could be assessed to fully understand the probability of failure during ESTO operation and how to cope with such a failure.
Once the need for the flap deflection schedule was established, a flap deflection schedule from 30° to 62° was iteratively optimized. Figure 11.1.2 presents a plot of the flap deflection
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angle, instantaneous velocity, and acceleration as a function of time during the ground roll for the Special Forces mission at hot and high conditions.
120 Velocity, ft/s Acceleration, ft/s2 100 Flap Angle, deg
80
60
40
20
0 0 1 2 3 4 5 6 7 8 9 10 time, sec
Figure 11.1.2 – Flap Deflection Schedule and Corresponding Velocity and Acceleration for Hot and High Conditions.
This plot shows that the flap angle was held constant for the first 2 seconds of the ground roll to achieve the maximum possible horizontal thrust and corresponding acceleration. From this point on, the flaps were deflected at a rate of 4.5°/sec to achieve maximum flap deflection of
62° and a lift coefficient of CL = 3.53 at takeoff. The rate of 4.5°/sec iteratively optimized to allow the C-328 flaps to reach the required 62° at the end of the ground roll while not exceeding the maximum rate of 9°/sec as specified by the Systems group.
Figure 11.1.3 presents the flap deflection schedule that was used for the Special Forces mission at sea level conditions.
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90
80 Velocity, ft/s Acceleration, ft/s2 70 Flap Angle, deg
60
50
40
30
20
10
0 0 0.5 1 1.5 2 2.5 3 3.5 4 time, sec
Figure 11.1.3 - Flap Deflection Schedule and Corresponding Velocity and Acceleration for Sea Level Conditions.
This plot shows that the flap angle was varied constantly at a rate of 8.2°/ sec throughout the ground roll. This constantly varying flap deflection schedule (no 30° deflection hold at the beginning of the ground roll) was mandated by the fact that the engines have significantly improved performance at sea level resulting in faster acceleration and a shorter ground roll. Without constantly varying flap deflection, the C-328 was unable to reach 62° during the ground roll without exceeding the maximum rate of 9°/sec.
For the AJACS mission, at both hot and high conditions and sea level conditions, the relaxed takeoff distance requirement allowed for less aggressive takeoff schedule than that employed for the Special Forces mission. DP is not required to achieve the desired 1500 ft takeoff distance. A constant flap deflection of 45°, employed throughout the ground roll, produced acceptable acceleration and a shorter ground roll than required by the mission parameters.
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11.1.2 Short Landing Ground Roll
The short landing ground roll presented the unique problem of developing a system capable of providing the required deceleration of the aircraft once contact with the ground was established. It was quickly decided that the complex landing routine required to land the aircraft in such a short field was beyond the capabilities of a human pilot, and therefore an automated landing system was implemented. A final optimized landing routine was created through an iterative process, where reverse thrust, distributed propulsion, brakes, spoilers and a drag parachute were all incorporated. Equation 11.2 presents the resulting expression for the instantaneous acceleration during the short landing ground roll.
11.2
Note that for analysis purposes, the drag chute was treated as a hemispherical shell and the flaps and spoilers were treated as flat plates with areas representative of the frontal area of the deflected flaps and spoilers.
Figure 11.1.4 presents a plot of the landing deceleration to illustrate the effects of the various systems employed during short landing operation. The same system was used for both sea level and hot and high conditions though the magnitude of the deceleration changed due to degraded engine performance at hot and high conditions.
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-11
-12 2 -13
-14
-15 Acceleration, ft/s
-16
-17 0 2 4 6 8 10 time, sec
Reverse Brakes applied Acceleration tapers off Thrust (0.8 seconds) as speed decreases
Spoiler deflection Drag pa rachute fully bleeds lift and deployed (2 seconds) increases drag (0.5 seconds)
Figure 11.1.4 – Landing Acceleration showing the Effect of Various Arresting Systems.
From this figure it can be seen that reverse thrust from the large turbofan engines is spooled up prior to touchdown to allow for full reverse thrust at touchdown. This capability has been proven by the C-179. Distributed propulsion was used during the final stages of decent as well as during approach to allow for a large value of CL which drives down the stall speed and consequently the touchdown speed (1.1 times the stall speed). At touchdown, the brakes, spoilers and parachutes are all deployed, but each system’s effects on the acceleration of the C- 328 is delayed to account for their individual deployment times. Half a second after touchdown the spoilers are fully deflected, 0.8 seconds after touchdown the brakes are fully applied and 2 seconds after touchdown the parachute is deployed. These times were chosen to account for the actuation time of the brakes and the deployment time of the spoilers and parachute.
Similar analysis was used for the AJACS mission; however, much like for the takeoff ground roll, the relaxed landing distance allowed for a less aggressive landing schedule than that
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employed for the Special Forces mission. Through iterative optimization, the landing deceleration proved acceptable without the use of a drag parachute. It is important to note that distributed propulsion was still employed during approach and flare to achieve the minimum touchdown speed possible. Additionally, full reverse thrust was still implemented throughout the ground roll.
11.2 Conventional Mission Segments Conventional mission segments (those not requiring distributed propulsion) were analyzed using methods available in the performance texts of Raymer39 and Torrenbeek49. This analysis provided the framework for the overall mission simulation and the method used remained unchanged for both missions.. The primary purpose of the conventional mission segment analysis was to generate total mission time, fuel burn results and mission segment flight speeds, which fed into other aspects of the design, such as fuel tank placement. The conventional mission segment analysis was identical for the AJACS and Special Forces missions with the exception of the cruise distance, payload weight, the additional idle segment for the Special Forces mission and loiter segment for the AJACS mission.
11.2.1 Unrestricted Takeoff and Landing
Unrestricted takeoff was achieved at 1.2 times the stall speed of the aircraft with no flap deflection. A screen height of 50 ft was used to compute the complete takeoff field length, including the ground roll and climb to the screen height. Unrestricted landing was achieved with approach at 1.2 times the stall speed, flare at 1.1 times the stall speed and touchdown at 1.1 times the stall speed. Brakes were the primary arresting mechanism during unrestricted landing as spoilers were not used. The field on which all take off and landings were performed was assumed to be that of soft turf, with a takeoff friction coefficient of 0.02 and landing friction coefficient of 0.2 as suggested by Raymer39. This is the worst case friction situation in austere conditions in which the aircraft is expected to operate.
11.2.2 Climb, Cruise and Descent
The climb distance was set as the difference between the cruising altitude and the takeoff altitude. As required by the mission specification, cruise was achieved at Mach 0.8 at 35,000 ft. The cruise distance was selected by taking the required combat radius of the aircraft (1000nm for the Special Forces mission and 500 nm for the AJACS mission) and subtracting the ground distance covered during the climb to altitude. As is standard for aircraft design, the ground
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distance covered during the descent was not included in the combat radius, as advised by Sam Wilson III, an industry expert and the “Right Reverend of the STOVL Faithful”. The cruise was essentially extended to the required landing field.
11.2.3 Loiter and Idle Segments
The Special Forces mission required a 1.5 hour idle segment after the first landing in order to remain combat ready. For this segment the Trent-895 engines were run at idle.
The AJACS mission required an additional loiter of 45 minutes prior to the final mission landing. This was modeled in the same manor as the cruise simulating the aircraft to hold at a near optimum altitude while awaiting clearance to land at a specific destination.
11.3 Mission Simulation A final performance code was created to analyze both missions incorporating the short takeoff and landing analysis as well as the conventional mission segment analysis. Starting at the unrestricted takeoff and progressing through the mission, fuel burn was calculated using fuel consumption data for both the DP and Trent-895 engines at all mission altitudes and throttle settings. This information was incorporated into the code to account for the change in weight of the aircraft throughout each mission. As a required factor of safety, an additional 10% of the mission total fuel required was incorporated as contingency fuel as well as enough fuel for an additional 45 minute cruise. Finally, 7,716 lbs was included as diversion fuel.
Table 11.3.1 presents the results of the mission simulation for both the AJACS mission and the Special Forces mission at both sea level and hot and high field conditions.
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Table 11.3.1 - Mission Analysis Summary
Short Takeoff Short Landing Total Mission Time Fuel Required
(ft) (ft) (hr) (lb)
Special Forces 178.0 298.0 6.57 131,993 Sea Level
Special Forces 515.4 500.7 6.37 131,078 Hot and High
AJACS 733.2 522.6 3.19 85,257 Sea Level
AJACS 883.2 616.9 3.03 76,794 Hot and High
From this table, it can be seen that the C-328 achieved the nominal short takeoff and landing requirements for the AJACS mission at both hot and high and sea level conditions. Additionally, with a shorter required cruise distance, the AJACS mission required less fuel and less mission time than the Special Forces mission.
The Special Forces mission proved to be much more challenging to meet the desired takeoff and landing requirements. An iterative optimization process was used to achieve a balanced ground roll distance for this mission at hot and high conditions as this is the primary mission for the C-328. As seen in Table 10.3.1, the C-328 was able to takeoff in 515.4 ft and land in 500.7 ft yielding a nearly balanced ground roll. Incorporating the landing gear stance of 55.1 ft, the C-328 is able to operate with a balanced short field landing strip of 571 ft. While this does not meet the nominal 328 ft field length, a balanced short field of 571 ft is significantly shorter than all other existing aircraft in its class.
11.4 Range and Endurance Figure 11.4.1 presents the payload vs. range diagram for the C-328.
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4 x 10
7 Range Envelope 66200lb max fuel range 6 26500lb max fuel range Max ferry range Absolute max ferry range 5
4
Payload (lb) 3
2
1
0 0 1000 2000 3000 4000 5000 6000 7000 Range (nm)
Figure 11.4.1 - Payload vs. Range Diagram
Carrying the AJACS mission payload of 66,200 lb, the C-328 has a max fuel range of 3,289 nm and a corresponding endurance of 7.13 hours. With the lighter Special Forces mission payload of 26,500 lb, that range is extended to 4,917.4 nm and the endurance is increased to 10.66 hours. If all payload is removed, the C-328 can achieve a maximum ferry range of 5510.1 nm. If all payload is removed and the payload is replaced with internal fuel stores, the C-328 can achieve an absolute maximum ferry range of 6179.4 nm.
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12. COST
12.1 Introduction to Aircraft Associated Costs In evaluating the cost of an aircraft, a quoted price may refer to a number of different costs that make up the sale of an aircraft. Thus, comparing costs of aircraft is futile unless the same type of costs for each aircraft is evaluated. Refer to Figure 12.1.1 for an illustrative guide to the breakdown of costs associated with the life cycle of an aircraft, defined as the total span of the design, production, maintenance, and disposal. The first major element represents costs associated with research, development, testing, and evaluation (RDT&E). This includes the technology research, design engineering, prototype fabrication, flight and ground testing, etc.
Figure 12.1.1 - Elements of Life Cycle Cost39
The second element is aircraft production cost, commonly known as the “flyaway” cost. This is the cost associated with any labor and material costs used to manufacture the aircraft, including airframe, engines, tooling, and aircraft systems. Program cost, which is frequently the quoted price for a new military aircraft, is the total cost to develop and deploy a new aircraft into the military inventory. This would be the RDT&E plus fly-away costs, as well as any costs
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associated with any special ground facilities required for deployment. The unit cost for an aircraft can be computed as the program cost per number of aircraft produced and supported.
As can be seen in Figure 12.1.1, Operations and Maintenance (O&M) costs usually make up the highest percentage of the life-cycle cost of an aircraft. This cost covers fuel, oil, aircrew, and various indirect costs. The last element of the life-cycle cost is that required to dispose of the aircraft. While it is frequently ignored in life cycle cost estimation, it commonly makes up 10% of the life cycle cost.39
12.2 Current Military Transport Market One of the major parameters not given in the mission specifications that greatly affected any conceivable cost model was the number of desired aircraft to be produced. In deciding upon this two issues had to be taken into account: 1. the number of competitor transport aircraft in service and 2. the cost benefits of producing more aircraft ( learning curve effect).
Table 12.2.1 summarizes performance characteristics of similar aircraft in service today. As can be seen in this table, the C-328 is extremely competitive in today’s modern military transport aircraft market, excelling in three of the four proposed performance measurement categories. Thus, except in carrying payload, the C-328 could perform the mission profiles of these transport aircraft. Table 12.2.1 also indicates the number of comparable aircraft that are in service for the proposed UK and US customers.
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Table 12.2.1 - Performance measurements and number of comparable aircraft
A400M C-130J C-17 C-328
Max TOGW (lb) 310,852 155,000 585,000 330,693
Max Payload 81,571 42,000 170,900 66,139 (lb)
Range w/ Max 1782 1800 2420 3289 Payload (n mi)
Min TO Dist (ft) 3084 5160 7600 178
Min Landing 2050 2020 2700 546 (ft)
# in UK Service 25 50 6 --
# in US Service 0 435 174 --
Unit Cost $140 million $62 million $202 million $151 million (Year) (2008) (2008) (2008) (2009)
Cost /lb. Wempty $909 / lb $1808 / lb $716 / lb $1020 / lb
The second element in analyzing the number of proposed aircraft to be produced is the cost-benefit analysis of producing more aircraft. Referred to as the “learning curve effect,” it is intuitive that the more aircraft of one specific type produced the cheaper it becomes. This is illustrated in Figure 12.2.1 below. Cost estimates provided by the DoD for the dynamic C-130J and C-17 programs are shown for comparison51, 52.
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With these two major factors taken into account, it could be concluded that the more aircraft produced the faster the unit price per aircraft would decrease. Table 12.2.1 further expanded that more C-328 aircraft could not only compete in the modern transport market, but could replace many aircraft in service with better performance requirements. Thus, to quantify this approach, it was decided that a cost model would be performed for production of 200 C-328 aircraft over a thirty year service term.
Estimated Unit RDT&E + Fly-Away Cost versus Production Quantity 1.4
1.2 DAPCA IV C-17 1 C-130J 0.8
0.6 away Unit Cost ($ Billion) ($ Cost Unit away - 0.4
0.2 RDT&E + Fly 0 0 50 100 150 200 250 300
Production Quantity
Figure 12.2.1 - Learning curve affect on cost: JFTL with comparable aircraft
12.3 Estimating RDT&E, Flyaway, and Unit Cost A public aircraft cost estimating relationship (CER) “DAPCA IV” was used in evaluating the cost of the JFTL program. DAPCA is a continuously updated set of CER’s for conceptual aircraft design developed by the RAND Corporation for the Development and Procurement Costs of Aircraft model. DAPCA simply estimates the hours required for RDT&E and production by the engineering, tooling, manufacturing, and quality control groups, which are multiplied by the appropriate hourly rates to yield costs. Development support, flight test, and manufacturing material costs are directly estimated by DAPCA. The model is highly dependent upon only three variables: aircraft empty weight, maximum cruise speed, and number of aircraft produced. Smaller costs such as engine costs, if unknown, are estimated based on turbine inlet and thrust.
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Using equations from Raymer39 the RDT&E and flyaway costs were evaluated for the JFTL program. A major unknown was the material factor, which refers to the “empirical factor” that accounts for materials other than aluminum, such as titanium. Since the provided C-328 design had similar titanium flap settings as the C-17, the DAPCA model was calibrated with this factor until it mirrored the advertised program cost of the C-17, within an error of 5%. This resulted in a 10% increase in manufacturing hours.
The program cost for producing 200 aircraft is listed in Figure 12.3.1 as $30.8 Billion. This figure also goes further in showing the percentage breakdown of program costs associated with tooling, propulsion, and any other items listed in Raymer39.
Figure 12.3.1 - Tabulated RDT&E + Fly-away Costs with Breakdown of Program Costs
Referring back to Table 12.2.1, the cost comparisons of similar aircraft are presented with these DAPCA model estimations. Aircraft costs are commonly adjusted to the same level by comparing cost per pound empty weight, as done in Table 12.2.1. While the C-328 out-performs each competitor aircraft in performance measurements, it fails to do so in a cost effective manner. This is due to the new propulsion technology, complexity of the system, and added weight of the titanium flaps.
Because it is more than common that the estimated weight of the aircraft will increase or decrease through design and production, it is important to show the sensitivity of the aircraft weight on cost. Since empty weight is a major parameter in Raymer’s program cost model, this relationship can easily be illustrated, as done in Figure 12.3.2 below.
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Estimated Total RDT&E and Fly-Away Cost versus Empty Weight 16.2 16 15.8 $56 / lb 15.6 15.4 15.2 DAPCA IV 15
away Unit Cost ($ Million) ($ Cost Unit away 14.8 Final Iteration 14.6 14.4 135000 140000 145000 150000 155000 160000 165000 170000 RDT&E + + - RDT&E Fly Empty Weight (lb)
Figure 12.3.2 - Unit RDT&E + Fly-away Costs Varying with Empty Weight for Producing 200 Aircraft
An estimated sensitivity of $56 dollars unit cost per pound is more than commonly accepted within the possible growth of any aircraft in production43.
12.4 Estimating Operation, Maintenance, and Disposal Costs For estimating O&M cost, a different CER model adapted from Dr. Jan Roskam4 was used. Roskam’s model breaks O&M cost into seven categories, listed in Figure 12.4.1. Major parameters of this model include: fuel weight, crew and labor rates, maintenance man hours per flight hour, and mission time.
The least cost beneficial aspect of the new DP system was its increased maintenance cost. Adding more engines and embedding them into the wing intuitively increases the amount of man hours servicing the aircraft between missions. To account for this, the maintenance man hours per flight hour was scaled up from a C-130J43, which is listed as using 22 hours, to 30 hours per flight hour. Using the equations listed in Roskam43, the O&M cost for servicing 200 aircraft is listed in Figure 12.4.1. This figure also shows the breakdown of program costs associated with fuel, personnel, and other items listed in Roskam43.
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Figure 12.4.1 - Operation and Maintenance Costs
The operational cost per hour of the C-5, C-17, and C-328 are shown in Figure 11.4.2 below. This figure attempts to show the C-328 has an operational cost slightly less than the C-17, and well below that of another military transport, the C-5. The C-17 has a much higher gross weight, which accounts for its higher cost at lower flight hours per year. The high thrust of the C-328 causes its O&M cost to rise drastically with flight hours per year because of its fuel burn increase with numerous engines.
25
20 C-17
15 C-5 JFTL
Millions 10
5 Annual O&M cost per per in Hour O&M cost Annual 0 0 500 1000 1500 Operation Hours per Year
Figure 12.4.2 - O&M Costs per year Varying with Operation Hours per Year for Transport Market
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Finally, the disposal cost was taken to be 10% of the total life-cycle cost, in accordance with Raymer’s cost model.39
12.5 Life Cycle Costs Thus, the life cycle costs for designing 200 aircraft that fly 2000 hours per year for a service of 30 years are summarized in Table 12.5.1 below.
Table 12.5.1 - Estimated JFTL Life-Cycle Costs of 200 Aircraft over 30 years service life
RDT&E + Fly-away Cost $30.84 Billion
Operations and Maintenance Cost $204.10 Billion
Disposal Cost $2.37 Billion
Total Program Life-Cycle Cost $237.31 Billion
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13. CONCLUSION The design of the C-328 Ostrich began with a harmonization of three individual programs of the Army, Air Force, and Special Forces into one mission profile, the JFTL program. The ESTOL requirement constrained to the distance of a full-sized soccer field was the driving force behind this design. To meet this extraordinary high-lift need, a new technology known as distributed propulsion was researched and implemented for blowing numerous engine exhausts across a wing. The idea of embedding these smaller engines into the wing was proposed to decrease skin friction drag and engine wash, thus increasing the aerodynamic efficiency of the distributed propulsion system. In order to house this system, as well as to increase lift during low speeds of takeoff and landing, a low aspect ratio, low wing loading, and relatively high wing span were chosen in the wing design. To meet the structural complexity of an embedded wing design, it was decided to have a straight trailing edge with rear spar, thus giving the aircraft a delta-wing shape. Next, a super critical airfoil was selected based on the transonic drag and stall characteristics experienced during the required Mach 0.8 cruise.
A major design decision in regard to this new propulsion system was the number of smaller engines desired. This was geometrically constrained because of the maximum wing span required to fit within a soccer field as well as the desired t/c of the airfoil to handle transonic speeds. Thirty-six Honda/GE HF-120 engines were chosen because of their size, relatively high thrust to weight, and previous experience in service today. Because the DP gross takeoff thrust did not meet the ESTOL requirements, two primary engines were added to provide the necessary takeoff thrust. The up and coming Rolls-Royce Trent 895 fit into this required thrust regime and was chosen for its fuel efficiency, low weight, and predicted cost benefits.
Because the cargo holding area dimensions were included in the JFTL requirements, the shape of the fuselage was designed to mold around the box-shaped bay. The size of the fuselage, however, was based on a compromise between the required short ground roll and the desired tail moment arm. For balance, weights were assigned to each component of the aircraft along with moment arms to compute the CG, which affected stability and control. As the design evolved, components were drawn and updated in models created with AutoCAD and SolidWorks.
Using the final numbers from the iterative design process, performance analysis was conducted to ensure the proposed C-328 met all requirements. Finally, the cost was estimated to
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explore if the proposed aircraft was competitive in the highly dynamic, medium sized transport market.
In response to the USAF, US Army, and Special Forces programs, the Virginia Tech/ Loughborough University International Aircraft Design team presents the C-328 Ostrich.
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References
1. "The HF120 Engine". GE Honda Aero Engine. General Electric.
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34. Newberry, C., and J. Wimpress. YC-14 STOL Prototype: It's Designm Development and Flight Test. Reston, VA: AIAA Education Series, 1998. 35. Niu, M. C. Y. Airframe Structural Design. Conmilit Press Ltd., 1995. 36. Norton. B. STOL Progenitors: The Technology Path to a Large STOL Aircraft and the C-17A. Reston, VA: AIAA Inc, 2002. 37. Pallett, E. H. J. Aircraft Electrical Systems. Third ed. Essex: Longman Scientific & Technical, 1987. 38. Pallett, E.H.J. Aircraft Instruments & Integrated Systems. 1st ed. Essex: Longman Scientific & Technical, 1992. 39. Raymer, Daniel P. Aircraft Design: A Conceptual Approach. Reston, VA: AIAA, 2006. 40. Red Hammer Consulting Limited. TORNADO Vortex Lattice Method. Computer Program. 2008. 41. Roskam, J. Airplane Design, Part VI, Preliminary Calculation of Aerodynamic, Thrust and Power Characteristics. Ottawa, Ka: Roskam Aviation and Engineering Corporation, 1989. 42. Roskam, J.. Airplane Design, Part VII, Flight Dynamics and Automatic Flight Controls. Ottawa, Ka: Roskam Aviation and Engineering Corporation, 1989. 43. Roskam, J. Airplane Design Part VIII: Airplane Cost Estimation: Design, Development, Manufacturing and Operating. Ottawa, KS: Roskam Aviation and Engineering Corporation, 1990. 44. Schmitt, Vernon R., James W. Morris, and Gavin D. Jenney. Fly-by-Wire A Historical and Design Perspective. 1st ed. Warrendale, PA: Society of Automotive Engineers, 1998. 45. Spence, D. A. "The Lift on a Thin Aerofoil with a Jet-Augmented Flap." The Aeronautical Quarterly August (1958). 46. Spitzer, Cary R. Digital Avionics Handbook. Avionics: Development and Implementation. 2nd ed. London: CRC Press, 2006. 47. Stinton, D. The Anatomy of the Aeroplane. 2nd ed. UK: Blackwell Science Ltd, 1998. 48. Tooley, Mike, and David Wyatt. Aircraft Communications and Navigation Systems. 1st ed. Oxford: Elsevier, 2007. 49. Torenbeek. E. Synethesis of Subsonic Airplane Design. Ed. Delft University Press. MA: Kluwer Academic Publishers, 1982. 50. "Trent 800. Civil Aerospace - Large Aircraft Engines." Rolls-Royce PLC. April 26, 2009
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