an international cooperative design effort between virginia tech...

An international cooperative design effort between

Virginia Tech and Loughborough University

presents:

May 12, 2009

JFTL – C-328 Ostrich Final Report

Senior Design Project 3 May 2009

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.

148.2124

.26

81.08

MAC POSITION

27.15

GENERAL ARRANGEMENT DRAWINGJFTLSCALE 1:150

10-APRIL-09 UNITS FEETISSUE 9

5 10 15 20 250METERS

20 40 60 800

FEET

AA

BB2.

89

50.5565.21

20.5

426

.31

LANDING GEARLAYOUT (REF)

JFTL

CC

SECTION A-ADISTRIBUTED PROPULSION TAKEOFF CONFIGURATION

SECTION B-BDISTRIBUTED PROPULSION TAKEOFF CONFIGURATION

SECTION B-BSPOILER UP AND FLAP DOWN CONFIGURATION

SECTION B-BCLEAN CONFIGURATION

SECTION A-ACLEAN CONFIGURATION

SECTION C-CCLEAN CONFIGURATION

SECTION C-CDISTRIBUTED PROPULSION TAKEOFF CONFIGURATION

VIEW SHOWINGTAKEOFF TAIL

CONFIGURATION

VIEW LOOKING DOWNSHOWING FUSELAGEFRAME LOCATIONS

1.83

4.85

DATA SUMMARYPERFORMANCE

MAX RANGE AT 12T PAYLOAD 4917 nmMAX RANGE AT 30T PAYLOAD 3289 nmMAX FERRY RANGE 5510 nmCRUISE MACH NUMBER 0.8CRUISE ALTITUDE 35,000 ft

POWER PLANT

MAIN ENGINES 2 X ROLLS-ROYCE TRENT 895DISTRIBUTED ENGINES 36 X HONDA/GE HF-120TRENT 895 THRUST 93,400 lbHONDA/GE HF-120 2,050 lb

WEIGHTS

OPERATIONAL WEIGHT EMPTY 143,141 lbMAX TAKE OFF WEIGHT 330,693 lb

DIMENSIONS

LENGTH 147.9 ftWING SPAN 177.4 ftWING AREA • • • • •• ••••••••••••••ASPECT RATIO 4.97WING LEADING EDGE SWEEP •••••••••••••••••• • •••••••••••••••••••••WHEEL BASE 40.38 ftWHEEL TRACK 26.31 ftTAIL SCRAPE ANGLE •••••••••••••••••• • •••••••••••••••••••••••

David BrindleyVictoria CopeScott FerryRyan HurrilBen King

Simon LangleyAlex McMillanChris SkinnerSeb Wilkes

Tyler AaronsDavid Gladson

Ryan MerittChris Olien

Grant ParrishWendy Pifer

Steve SikorskiShadie Tanious

VT MEMBERS LU MEMBERS

INTERNATIONAL DESIGN TEAM

2.51

177.

49

110.

14

9.86

97.7

761

.62

10.17

KEY

DISTRIBUTED PROPULSION ENGINE

CONTROL SURFACE

CENTER OF GRAVITY



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



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



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



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



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



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 - Dutch Roll Undamped Natural Frequency, ωnD ................................................................................ 64

Table 9.4.4 - Dutch Roll Damping Ratio, ................................................................................................................... 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



Nomenclature English AC – Aerodynamic center AOA – Angle of attack AR – Aspect ratio b - Span c - Chord c = Wing mean chord c/cf

- Lift Coefficient CG – Center of gravity c

– Wing chord to flap chord ratio

ht = Horizontal tail volume coefficients cvt = Vertical tail volume coefficient DP – Distributed propulsion F – Lift correction for aspect ratio Kg- Gust correction factor L - Lift L/D – Lift to drag ratio Lht = Horizontal tail moment arm Lvt = Vertical tail moment arm M – Mach Number mjvj – Jet-momentum S – Planform Area S’ – Wing blown area sfc – Specific fuel consumption Sht = Horizontal tail area Svt

– Planform exposed area = Vertical tail area

– Planform wetter area t – Thickness t/c – Thickness to chord ratio T/W – Thrust to Weight ratio TOGW – Takeoff Gross Weight V - Velocity

- Blowing coefficient ∂C

Greek

l/∂δ –Lift curve slope α – Airfoil incidence β = Sideslip angle δ – Flap deflection angle λ – Part span of jet flaps Λ - Sweep ν – Fuselage cut-out area ρ - Density

Acronyms ACS – AirCraft Synthesis AJACS - Advanced Joint Air Combat System Program ANSYS – AirCraft Synthesis AOE – Aerospace and Ocean Engineering AVID – Air Vehicle Integrated Design CER – Cost Estimating Relationship CFD – Computational Fluid Dynamics CVO – Chief Visionary Officer DAPCA – Development and Procurement Cost of Aircraft DARPA – Defense Advanced Research Projects Agency DoD – Department of Defense ESTOL – Extreme Short Takeoff and Landing FAR – Federal Aviation Regulations FEM – Finite Element Model FOD – Foreign Object Debris ICS – Initial Concept Sketching JFT – Jet Flap Theory JFTL – Joint Future Theatre Lift Program JHL – Joint Heavy Lift Program LU – Loughborough University MAC – Mean Aerodynamic Chord NACA – National Advisory Committee on

Aeronautics O&M – Operations and Maintenance OEI – One Engine Inoperative OWE – Operating Weight Empty RDT&E – Research, Development, Testing, and

Evaluation STOL – Short Takeoff and Landing STP – Standard Temperature and Pressure USAF – United States Air Force USB – Upper Surface Blowing VT – Virginia Tech VTOL – Vertical Takeoff and Landing

Units °F = Fahrenheit ft = Feet hr = Hour in = Inches kts = Knots lb = Pounds sec = seconds



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:



Tyler Aarons – Mission/Performance and Report Coordinator

David Gladson – Structures

Ryan Meritt – Propulsion

Chris Olien – Aerodynamics and Configuration/CAD

Grant Parrish – Weights

Wendy Pifer – Stability and Control

Steve Sikorski – Cost/Economics and Configuration/CAD

Shadie Tanious – Team Leader and Systems



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

Payload: 57,300 lb

- Stryker

66,200 lb

- 7 x 463 L pallets

26,400 lb

- 2 HMMWVs + 12 man crew

Cruise Speed: Mach 0.4 Mach 0.8 Mach 0.8



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.

Figure 2.2.1 - Special Forces Mission Schematic

Figure 2.2.2 – AJACS Mission Schematic

Unrestricted TO / LD 328 ft ground roll

1.5 hour idle

1000nm cruise, M=0.8, 35,000 ft

SPECIAL FORCES 26,500 lb payload

Unrestricted TO / LD 1,500 ft ground roll

500nm cruise, M=0.8, 35,000 ft

AJACS 66,200 lb payload

45 minute loiter



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

(Takeoff and Landing) 1500 ft 328 ft

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.



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.

The final row represents the optimized C-328 Ostrich.



Row 1: Initial Concepts

Row 2: Joint Concepts

Row 3: Down Selection #1

Row 4: Down Selection #2

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.



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

4.1.1 Wing Geometry

equation is the wing area, S, which was returned to the user as an output.

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



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

4.1.3 Center of Gravity

. 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.

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



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



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 Shape

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.

10



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



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)



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



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 Inlet

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,

6



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.



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 Design

The primary reason for selecting this unconventional configuration was the increase in lift

offered relative to an un-blown wing. Research conducted by AVID

3

3 explored the effects of using

distributed propulsion in conjunction with Upper Surface Blowing (USB); the results of the study



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 Distribution

6.2 Jet Flap Theory and Integration

3

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.

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 ( ).



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), part span jet flaps (λ) and the fuselage cut-out area (ν). Incorporating all of

these presents an accurate estimation of the lift coefficient which is tailored 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



downwards. The benefit is greater lift potential from the wing, as the lift curve slope (∂Cl/∂δ) is

steeper than that of a traditional jet flap as shown by Reference45,56

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 sizes

.

45,56

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.

. A larger ratio

would result in ground interference at full flap deflection, while any smaller ratio would reduce lift

potential from 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

6.2.3 Stage Flap Deflection

,

the induced pitching moment was found to shift 7.7% mean aerodynamic chord (MAC) rearward.

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



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.



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.



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.

Weight (lb) 390

Fan Diameter (in) 16

Bypass Ratio 2.9

Pressure Ratio 24

Cruise SFC (lb/lbf-hr) 0.6 Figure 6.3.2 - Engine Characteristics1

Table 6.3.2 - HF120 Engine Characteristics1



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

PERFORMANCE Sea-Level Condition Hot & High Condition

C 1.75 μ 1.00

Engine No: 36 36

CL 4.21 wing 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.



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 Comparison

50

Weight (%) Two Engines

(PW4090) Four Engines

(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 $/year)

10% 72.5 79.6

Static Thrust-to-Net Weight Ratio 40% 5.38 4.39

Total Cruise SFC (lb/lbf-hr) 25% 1.15 1.36



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 Engines

50

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



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.

Figure 6.4.2 - Rolls Royce Trent 895 Engine

6.4.2 Nacelle Design

50

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.

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

Table 6.4.3 - Rolls Royce Trent 895 Engine

Characteristics50



Figure 6.4.3- Long Duct Nacelle

The dimensions of the nacelle were calculated using the methods found in Johnson

22

22

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.

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

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

Table 6.4.4 - Nacelle Dimensions



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 Comparison

6.4.4 Positioning

22

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.



Figure 6.4.6 - Horizontal Engine Placement

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.

23

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



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



Figure 6.4.9 - Trent 895 SFC Thrust Curves for Hot and High and Sea Level Conditions



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.

:



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.

-1.5

-1

-0.5

0

0.5

1

1.5

2

2.5

3

3.5

0 100 200 300 400 500 600 700 800

Load

Fac

tor,

n

Equivalent Airspeed, knots

Vcruise = 462 knots

Vdive = 693 knots

Vstall = 222 knots

Vneg stall = 492 knots

66 fps

50 fps

25 fps

Flaps Down

nmax = 3.2



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 Table 7.2.1. The resulting values for this aircraft can be found in .



Table 7.2.1 - Web Thicknesses at Root and Tip for Wing Box Spars

FRONT in

CENTER in

REAR 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.

Figure 7.2.2 – Wing Box Layout

DP Engines

FLAP

STRUCTURE

Aileron FRONT SPAR

MID SPAR

REAR SPAR



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.

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.

DP Engine Ducting

Front Spar Mid Spar Rear Spar

Flap Retracted

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

Front Spar Mid Spar Rear Spar

Flap Retracted

Flap Deployed

DP Engine Ducting



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



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

11

. 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 . 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



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.

Figure 7.4.1 – Material Weight Breakdown

Aluminum Alloys50%

Titanium Alloy18%

Graphite/Epoxy20%

Steel7%

Other5%



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 data

.

39

Table 8.1.1 - Empirical Data for Transport Aircraft for Approximate Empty Weight Buildup

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.

Item

39

lb/ft Multiplier 2

Wing 10.0 S

Horizontal Tail

exposed planform

5.5 S

Vertical Tail

exposed planform

5.5 S

Fuselage

exposed planform

5.0 S

Landing Gear

wetted area

.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

.



method was implemented to achieve this. In lieu of custom regression analysis, equations provided

in Raymer39 Table 8.1.2 were utilized. The results of this analysis are shown in .

Table 8.1.2 – Component Group Weights and Moments

Groups

Weight x-bar x-Moment 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 Stryker Mission 214461 1385 11137531

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



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



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.

Figure 8.2.2 - Potato Plot showing the CG Envelope of the AJACS Mission.

Figure 8.2.3 Potato Plot showing the CG envelope of the Special Forces Mission

Stryker Mission

130000

150000

170000

190000

210000

230000

250000

270000

290000

310000

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)

Wei

ght (

lbs) Varying Fuel, Full Payload

Varying Payload, Full Fuel

Varying Fuel, No Payload

Varying Payload, No Fuel

Special Forces Mission

100000

150000

200000

250000

300000

350000

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)

Wei

ght (

lbs) Varying Fuel,Full Payload

Varying Payload, Full Fuel

Varying Fuel, No Payload

Varying Payload, No Fuel

AJACS Mission



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.



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

Table 9.2.1 – Summary of Control Surface Sizing

.

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.

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



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

Table 9.3.1

. The moments of inertia are

shown in .

Table 9.3.1 – Mass Moments of Inertia

Mxx (slugs/ft2 Myy (slugs/ft) 2 Mzz (slugs/ft) 2

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



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 Requirement 0.35 – 1.30 0.30 – 2.00

Calculated 3.43 0.21

Table 9.3.3 - Phugoid Minimum Damping Ratio

Phase C Phase B

Level I MIL Requirement 0.04 0.04


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.



Table 9.4.1 - Lateral Directional Stability and Control Derivatives

Takeoff Cruise

C -0.76 yβ -0.79

Cl -0.36 β -0.36

C 1.40 nβ 1.41

C 0.0 yda 0.0

C 0.046 lda 0.046

C 0.0064 nda 0.0064

C 0.016 ldr 0.017

C 0.085 ndr 0.91

C 0.78 yr 0.78

C -0.55 nr -0.55

C 0.12 lr 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 Requirement 12 sec 20 sec

Calculated 11.42 sec 0.28 sec



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

Table 9.4.3 - Dutch Roll Undamped Natural Frequency, ω

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.

nD

Phase C Phase B

Level I MIL Requirement 0.4 sec 0.4 sec


Table 9.4.4 - Dutch Roll Damping Ratio,

Phase C Phase B

Level I MIL Requirement 0.08 0.08


Table 9.4.5 - Dutch Roll Real Root Part Value,

Phase C Phase B

Level I MIL Requirement 0.10 rad/sec 0.15 rad/sec

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



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 Required 1.4 sec 1.4 sec




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 Gear

10.1.1 Layout and Arrangement

57

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.



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.

(Dimensions shown in ft)



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



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

Figure 10.1.5

, the loading on

each tire has been calculated as shown in below.

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.

(Loadings in lb)



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

10.1.4 Special Features

. These calculations indicate that the

strut length for the landing gear should be at least 25 inches in all cases.

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



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,



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.



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.

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

W

Tmain

TDP D Ffriction

L



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

Once the need for the flap deflection schedule was established, a flap deflection schedule

from 30° to 62° was iteratively optimized.

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.

Figure 11.1.2 presents a plot of the flap deflection



angle, instantaneous velocity, and acceleration as a function of time during the ground roll for the

Special Forces mission at hot and high conditions.

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

Figure 11.1.3

= 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.

presents the flap deflection schedule that was used for the Special Forces

mission at sea level conditions.

0 1 2 3 4 5 6 7 8 9 100

20

40

60

80

100

120

time, sec

Velocity, ft/s

Acceleration, ft/s2

Flap Angle, deg



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.

0 0.5 1 1.5 2 2.5 3 3.5 40

10

20

30

40

50

60

70

80

90

time, sec

Velocity, ft/s

Acceleration, ft/s2

Flap Angle, deg



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.



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

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

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.

0 2 4 6 8 10-17

-16

-15

-14

-13

-12

-11

time, sec

Acc

eler

atio

n, ft

/s2

Spoiler deflection bleeds lift and increases drag (0.5 seconds)

Brakes applied (0.8 seconds)

Drag parachute fully deployed (2 seconds)

Reverse Thrust

Acceleration tapers off as speed decreases



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

11.2.1 Unrestricted Takeoff and Landing

. 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.

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

11.2.2 Climb, Cruise and Descent

. This is the worst case friction situation in austere

conditions in which the aircraft is expected to operate.

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



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.



Table 11.3.1 - Mission Analysis Summary

Short Takeoff (ft)

Short Landing (ft)

Total Mission Time (hr)

Fuel Required (lb)

Special Forces Sea Level 178.0 298.0 6.57 131,993

Special Forces Hot and High 515.4 500.7 6.37 131,078

AJACS Sea Level 733.2 522.6 3.19 85,257

AJACS Hot and High 883.2 616.9 3.03 76,794

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.



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.

0 1000 2000 3000 4000 5000 6000 70000

1

2

3

4

5

6

7

x 104

Range (nm)

Pay

load

(lb)

Range Envelope66200lb max fuel range26500lb max fuel rangeMax ferry rangeAbsolute max ferry range



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 Cost

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

39



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.



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

(lb) 81,571 42,000 170,900 66,139

Range w/ Max

Payload (n mi) 1782 1800 2420 3289

Min TO Dist (ft) 3084 5160 7600 178

Min Landing

(ft) 2050 2020 2700 546

# in UK Service 25 50 6 --

# in US Service 0 435 174 --

Unit Cost

(Year)

$140 million

(2008)

$62 million

(2008)

$202 million

(2008)

$151 million

(2009)

Cost /lb. W $909 / lb empty $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.



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.

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.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

0 50 100 150 200 250 300

RDT&

E +

Fly-

away

Uni

t Cos

t ($

Billi

on)

Production Quantity

Estimated Unit RDT&E + Fly-Away Cost versus Production Quantity

DAPCA IV

C-17

C-130J



Using equations from Raymer39

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 Raymer

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.

39

.

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.



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

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-130J

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.

43, 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.

14.4

14.6

14.8

15

15.2

15.4

15.6

15.8

16

16.2

135000 140000 145000 150000 155000 160000 165000 170000RDT&

E +

Fly-

away

Uni

t Co

st ($

Mill

ion)

Empty Weight (lb)

Estimated Total RDT&E and Fly-Away Cost versus Empty Weight

DAPCA IV

Final Iteration

$56 / lb



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.

Figure 12.4.2 - O&M Costs per year Varying with Operation Hours per Year for Transport Market

0

5

10

15

20

25

0 500 1000 1500

Ann

ual O

&M

cos

t per

Hou

r in

Mill

ions

Operation Hours per Year

C-17C-5JFTL



Finally, the disposal cost was taken to be 10% of the total life-cycle cost, in accordance with

Raymer’s cost model.

12.5 Life Cycle Costs

39

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



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



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.

JTFL – C-328 Ostrich Final Report


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