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PSU & Technion – International Collaboration
Two Seat Turbine Helicopter In Response to the 2006 Annual AHS International
Student Design Competition – Undergraduate Category June 2, 2006
ii
TABLE OF CONTENTS
LIST OF FIGURES..............................................................................................................................v
LIST OF TABLES ..............................................................................................................................vi
ABBREVIATIONS AND NOMENCLATURE................................................................................vii
Special Thanks ....................................................................................................................................ix
Proposal Requirements Matrix.............................................................................................................x
Introduction ..........................................................................................................................................1
Table of Physical Data .........................................................................................................................1
GrassChopper General Arrangement ...................................................................................................2
GrassChopper Structural and Major Systems Layout – Inboard Profile..............................................3
Section 1 – Aircraft Trade Study..........................................................................................................4 1.1 – GW estimation.........................................................................................................................4 1.2 – RP estimation ..........................................................................................................................4 1.3 – MDL and MRD estimation .....................................................................................................4 1.4 – TRD Estimation.......................................................................................................................4 1.5 – Angular Velocities Estimation ................................................................................................4
Section 2 – Airframe Design................................................................................................................5 2.1 – Design Criteria ........................................................................................................................5 2.2 – Material ..................................................................................................................................5 2.3 – Major Features.........................................................................................................................5 2.4 – Structural Floor Design ...........................................................................................................5 2.5 – Subfloor Design.......................................................................................................................5 2.6 – Landing Gear Design ..............................................................................................................6 2.7 – Tail Boom................................................................................................................................6 2.8 – Aft Fuselage Layout ................................................................................................................7 2.9 – Major Drive System Supports .................................................................................................7
2.9.1 – Static Mast Supports.........................................................................................................8 2.9.2 – Engine Deck .....................................................................................................................8
2.10 – Bulkheads ..............................................................................................................................8 2.11 – Cabin Structure......................................................................................................................8 2.12 – Firewalls ................................................................................................................................8
Section 3 – Turboshaft Engine Design.................................................................................................9 3.1 – Turboshaft Engine: Introduction .............................................................................................9 3.2 – PSU250 Operating and Performance Specifications...............................................................9 3.3 – PSU250 Section View...........................................................................................................10 3.5 - Intake......................................................................................................................................11 3.6 - Compressor ............................................................................................................................11
iii
3.7 - Burner ....................................................................................................................................12 3.8 - Turbine...................................................................................................................................12 3.9 - Exhaust...................................................................................................................................12 3.10 - Engine Performance Analysis – MATLAB Program .........................................................12 3.11 – Engine Cost Discussion.......................................................................................................15 3.12 – Operation and Summary......................................................................................................15
Section 4 – Gearbox Design...............................................................................................................16 4.1 – Planetary Gearbox .................................................................................................................16 4.2 – Main Rotor Gearbox..............................................................................................................17 4.3 – Tail Rotor Gearbox................................................................................................................17 4.4 – Static Mast.............................................................................................................................17 4.5 – Drive System Schematic .......................................................................................................18
Section 5 – Rotor Design ...................................................................................................................19 5.1 Main Rotor ...............................................................................................................................19
5.1.1 Theoretical Point of View .................................................................................................19 5.1.1.1. Hovering....................................................................................................................19 5.1.1.2. Autorotation ..............................................................................................................19
5.1.2 Number of Blades..............................................................................................................19 5.1.3 Main Rotor Chord .............................................................................................................19 5.1.4 Main Rotor Diameter ........................................................................................................19 5.1.5 Airfoil Section ...................................................................................................................20 5.1.6 Blade Twist .......................................................................................................................20 5.1.7 Taper..................................................................................................................................21 5.1.8 RPM ..................................................................................................................................21 5.1.9 Blade Structural Design ....................................................................................................21
5.2 Hub Design...............................................................................................................................22 5.3 Anti-Torque System .................................................................................................................22
5.3.1 The Choice of Anti-Torque System ..................................................................................22 5.3.2 Tail Rotor Design ..............................................................................................................22
Section 6 – Systems............................................................................................................................22 6.1 – Fuel System Design...............................................................................................................22 6.2 – Oil System Design.................................................................................................................23 6.3 – Cockpit Control Panel ...........................................................................................................24 6.4 – Crashworthy Seats.................................................................................................................24 6.5 – Flight Control System ...........................................................................................................24 6.6 – Active Tail Buffet Damping..................................................................................................25 6.7 – Landing Gear Systems ..........................................................................................................25 6.8 – Available Upgrades ...............................................................................................................25
Section 7 – Performance Analysis .....................................................................................................26 7.1 – Download-Force Estimation .................................................................................................26 7.2 Trim Analysis ...........................................................................................................................26 7.3 Performance .............................................................................................................................27
Section 8 – Manufacturing .................................................................................................................29
iv
8.1 – Manufacturing Techniques....................................................................................................29 8.1.1 – Electron-Beam Curing....................................................................................................29 8.1.2 – Paint-less Finish .............................................................................................................29 8.1.3 – Tool-less Assembly ........................................................................................................29
8.2 – Component Materials and Concepts .....................................................................................30 8.2.1 – Rotor Blades Materials and Manufacturing ...................................................................30 8.2.2 Airframe Materials and Manufacturing.............................................................................30 8.2.3 – Composite Airframe.......................................................................................................31 8.2.4 – Foam Sub-floor ..............................................................................................................31 8.2.5 – Integrated Ceramic Composite Firewall ........................................................................31 8.2.6 – Advanced Composite Joint Concept ..............................................................................31 8.2.7 – Extruded Aluminum Tail Boom.....................................................................................32 8.2.8 – Summary ........................................................................................................................32
Section 9 – Weights and Center of Gravity Location ........................................................................32 9.1 Weight Estimation....................................................................................................................32 9.2 C.G. Estimation ........................................................................................................................33
Section 10 – Cost Analysis.................................................................................................................33 10.1 – Description and Validation of Cost Model .........................................................................33 10.2 – Recurring Cost Breakdown .................................................................................................34 10.3 – Cost Record .........................................................................................................................34 10.4 – Direct Operating Cost..........................................................................................................35
Appendix A – MIL-STD-1374 Weight Statement .............................................................................38
v
LIST OF FIGURES Figure 1.1: GW versus Payload. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Figure 1.2: Power versus GW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Figure 2.2: Drive System Layout Iteration. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .7
Figure 3.1: PSU250 Section View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .10
Figure 3.2: PSU250 Exploded View. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
Figure 3.3: PSU250 – SFC versus SHP and Altitude, ISA, Jet-A fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
Figure 3.4: PSU250 – SFC versus SHP and Altitude, ISA+20C, Jet-A fuel. . . . . . . . . . . . . . . . . . . . . . . . . 14
Figure 3.5: Maximum SHP versus Altitude, 1300K max cycle temperature. . . . . . . . . . . . . . . . . . . . . . . . . 14
Figure 3.6: Power to weight versus SFC. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Figure 3.7: SFC versus SHP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
Figure 3.8: Engine Cost Comparison. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15
Figure 4.1: Static Mast. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
Figure 4.2: Drive System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Figure 5.1: Required Power for Hovering versus MRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
Figure 5.2: Relative Performance Improvement versus MRD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
Figure 5.3: Cd/Cl^(3/2) for Checked Airfoils. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
Figure 5.4: Required Power versus 1twθ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . 20
Figure 5.5: AOA Distribution along Main Rotor Blade . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . 20
Figure 5.6: Modeled Rotor Blade . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . 21
Figure 5.7: Final Rotor Blade Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
Figure 6.1: Oil System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23
Figure 7.1: Power in Forward Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
Figure 7.2: Command Angles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
Figure 7.3: Payload versus Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Figure 7.4: HOGE versus altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Figure 7.5: Max. Velocity versus Altitude. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
Figure 9.1: CG Location. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . 33
Figure 10.1: Composite GrassChopper Cost Breakdown. . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . 34
Figure 10.2: Record of the cost for the GrassChopper. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35
Figure 10.3: Comparison of leading piston engine trainers to GrassChopper models. . . . . . . . . . . . . . . . . . 35
vi
LIST OF TABLES
Table 2.1: Landing Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Table 3.1: PSU250 Operating Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
Table 3.2: PSU250 Performance Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
Table 3.3: Summary of cost related features for the PSU250 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15
Table 7.1: RFP Mission Phases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
Table 7.2: GrassChopper Performance compared the to Robinson R22 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
Table 8.1: Rotor Blade Materials Comparison. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Table 8.2: Airframe Materials Comparison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Table 8.3: Manufacturing Techniques and Materials Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
Table 10.1: Recurring Cost Breakdown. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
vii
ABBREVIATIONS AND NOMENCLATURE
Ab – Airfoil Section Area
AR – Aspect Ratio
Cd – Two Dimensional Drag Coefficient
CF – Centrifugal Force
CG – Center of Gravity
Cl – Two Dimensional Lift Coefficient
Cl3/4 – Representative Two Dimensional Lift Coefficient of Blade Section at ¾ of Span
Cm – Two Dimensional Pitching Moment Coefficient
CP – Continuous Power
E – Young Modulus
EGT – Exhaust Gas Temperature
FF – Forward Flight
GW – Gross Weight
HOGE – Hover Out of Ground Effect
I – Blade Rotational Moment of Inertia
Ib – Airfoil Section Bending Monet of Inertia
M – Rotor Blade Mass
m – Rotor Blade Mass per Foot
Mb – Bending Moment of Inertia
MCP – Maximum Continuous Power
MDL – Main Disc Loading
MR – Main Rotor
MRC – Main Rotor Chord
MRD – Main Rotor Diameter
Nb – Number of Blades
OAT – Outside Air Temperature
PL – Payload
r – Rotor Radius
RP – Required Power
RPI – Relative Performance Improvement
SFC – Specific Fuel Consumption
SHP – Shaft Horsepower
viii
Sy – Yield Strength
T – Rotor Thrust
TIT – Turbine Inlet Temperature
TR – Tail Rotor
TRD – Tail Rotor Diameter
Vd – Rate of Descent
Vi – Induced Velocity
Φf – Body Roll Angle
λ - Dimensionless Inflow
Θf – Body Pitch Angle 1twθ – Blade Twist (Washout)
θ0 – Collective Command Angle
θ1c – Lateral Command Angle
θ1s – Longitudinal Command Angle
σb – Bending Stress
σt – Tension Stress
Ω – Main Rotor Rotational Velocity
ΩT – Tail Rotor Rotational Velocity
δ − Main Rotor tip displacement
ix
Special Thanks Mr. Avi Attias MATA factory
Abdul H. Aziz The Pennsylvania State University
Dr. Robert Bill The Pennsylvania State University
Ms. Lynn Byers The Pennsylvania State University
Dr. Cengiz Camci The Pennsylvania State University
Peter Cipollo The Pennsylvania State University
Ms. Camala Daly Alumnus, The Pennsylvania State University
Mr. Chen Friedman Technion IIT
Mr. Jason Girven Alumnus, The Pennsylvania State University
Dr. Boris Glezer Solar Turbines/Optimized Turbine Solutions
Dr. Joe Horn The Pennsylvania State University
Dr. Gil Iosilevskii Technion IIT
Dr. Jun-Sik Kim The Pennsylvania State University
Dr. Vladimir Khromov Technion IIT
William Kong The Pennsylvania State University
Dr. George Lesieutre The Pennsylvania State University
Mr. Chuck Nearhoof Innodyn Turbines
Mr. Edmond Pope Innodyn Turbines
Prof. Omri Rand Technion IIT
Dr. Daniella Raveh Technion IIT
Prof. Aviv Rosen Technion IIT
Zihni Saribay The Pennsylvania State University
Dr. Edward Smith The Pennsylvania State University
Michael Smith Bell Helicopters
Mr. Ian Stock Schweizer Aircraft Corp
Joseph Szefi The Pennsylvania State University
Mr. Matthew Tarascio Sikorski Aircraft Corp
x
Proposal Requirements Matrix
Design requirement GrassChopper capability Sec./Page Two-place training helicopter Two-place helicopter capable of
pilot training Throughout
Conceptual, low cost, single turbine engine Inexpensive turboshaft engine design Section 2
Cost competitive Design with specific attention to cost and manufacturing Throughout
Must maintain normal standards of safety and reliability
Meets FAR 27 and MIL-STD requirements Throughout
Must include operating environment and characteristics that are important to a training
helicopter
Easily Visible Standard flight systems included; Rotor designed for good handling characteristics
Sections 3 and 6
Innovative manufacturing cost reduction Simple design, using composite materials, modular assembly
Section 3, Section 8
Cost analysis Assuming a production rate of 300 aircraft per year for 10 years Section 10
Good autorotative capability Maximized Main Rotor Diameter for Minimal Disc Loading Section 5
Enough fuel to hover for two hours at 6000ft (HOGE) on an ISA+20ºC day, carrying two 90
kg people and 20 kg of equipment
Fuel capacity of 41.4 gal (135.8 L), in dual tanks
Table of Physical
Data Performance, in general, should be superior to
current piston trainers Top speed of 115 knots, cruise
speed of 89 Knots. Maximum range 290 nm, maximum endurance 230
min @ 60 knots
Section 7.3
Proposal requirement GrassChopper proposal Sec./Page Table of Physical Data Included Page 1
MIL-STD-1374 Weight Statement Included Appendix A Recurring Cost Breakdown Included Section 10.2
Direct Operating Cost Breakdown Included Section 10.4 Performance Charts:
HOGE altitude vs. gross weight Payload vs. range
Altitude vs. maximum continuous speed
Included Page 28
Drawings: General Arrangement
Inboard Profile Engine Centerline Drawing
Drive System Schematic
Included
Page 2 Page 3
Page 10 Page 18
Description of configuration process selection Included Throughout Attention to proposed manufacturing process Included Section 8,
Throughout
1
Introduction
This report is the result of a successful collaboration between undergraduate students from two
Aerospace Facilities: one from The Pennsylvania State University, and the other from The Technion-Israel
Institute of Technology. The GrassChopper, presented here, is a two seat trainer helicopter powered by an
innovative turbine engine design, in response to the 2006 annual AHS student design competition. This
collaboration benefited the students by increasing the participants’ academic knowledge and design aspect as
well as their ability to work in teams coordinated far apart, which is becoming typical in today’s networking
environment.
Some of the main requirements are two hours HOGE capability at 6,000ft altitude on an ISA+20o
day, and maintaining a low cost design using innovative manufacturing processes. A new turbine engine
design and a comprehensive cost analysis was completed by the Penn-State team, while rotor design, outer
shape, and performance analysis were made by the Technion team, along with the cockpit internal layout.
Airframe design and manufacturing details was a combined effort.
Special attention was devoted to keeping the cost of the GrassChopper as low as possible while
maintaining standards of safety, reliability, and performance. Since this is a trainer helicopter,
crashworthiness was a central design criterion. The helicopter features a conventional tail rotor anti-torque
system, a new, cost-effective 250 hp turboshaft engine design, and a two bladed high efficiency main rotor in
a simple seesaw configuration. Performance analysis yielded a top velocity of 115 Knots, 290 miles
maximum range, and almost four hours (230 minutes) of endurance at a forward flight speed of 60 knots.
The final design is competitive with existing piston helicopters at a cost that is well under $300 thousand
dollars.
Table of Physical Data
Major Dimensions Main Rotor Diameter Tail Rotor Diameter
Main Rotor Chord Total Width
Overall Length Total Height
*see included drawings
26.0 ft (8.0 m) 4.0 ft (1.2 m) 0.6 ft (0.2 m) 6.7 ft (2.0 m) 30.4 ft (9.3 m) 9.0 ft (2.7 m)
Engine Power Take Off
Maximum Continuous
225.0 hp (167.7 kW) 202.5 hp (151.0 kW)
Weights Gross
Empty Useful load
1540.0 lb (700.0 kg) 806.0 lb (365.6 kg) 734 lb (333.0 kg)
Transmission Ratings Main Gearbox
Maximum Maximum Continuous
275 hp (205 kW) 220 hp (164 kW)
Fuel Capacity Tank One Tank Two
20.7 gal (78.3 L) 20.7 gal (78.3 L)
Tail Gearbox Maximum
Maximum Continuous
25 hp (18.6 kW) 20 hp (14.9 kW)
2
GrassChopper General Arrangement*
*Dimensions in feet
3
GrassChopper Structural and Major Systems Layout – Inboard Profile
NO. Description 17 Firewalls 18 Overhead I-Beams 19 Engine Deck 20 Side Supports 21 Structural Floor 22 Foam Subfloor 23 Lift Frames 24 Tail Boom Support Struts 25 Extruded Aluminum Tail Boom 26 Tail Rotor Drive Shaft Cover 27 Engine Mount 28 Rear Bulkhead 29 Landing Skids St
ruct
ural
Lay
out
NO. Description 1 Main Rotor Blades 2 Main Rotor Swash Plate Arrangement 3 Static Mast 4 Static Mast Support 5 Gear Box 6 Crashworthy Seats 7 Control Stick 8 Control Panel 9 Pedals 10 Collective Stick 11 Fuel Tank 12 Tail Rotor Drive Shaft 13 Horizontal Stabilizer 14 Tail Rotor 15 Landing Gear Dampers 16 Engine
Syst
ems L
ayou
t
4
Section 1 – Aircraft Trade Study
Figure 1.1: GW versus Payload 200 400 600 800 1000500
1000
1500
2000
2500
3000
3500
4000N=26Wlinear=3.41*PL+45.9
Gross Weight Vs. Payload
Payload[lb]
Gro
ss W
eigh
t[lb]
Hellicopter DataLinear FittingQuadratic Fitting
One of the most important stages of the preliminary design process consists of the initial sizing of the
vehicle. The following trade studies included parameters such as: RP, GW, MRD/MDL, Ω and ΩT, TRD. In
the present study three databases were used, these are the RAPID/RaTE package ([RAPI00]), JANES, and
Manufacturers’ official sites. A similar study ([Rand02]) was also used for the initial sizing. However, since
a study on lightweight helicopters was needed, correlations from
[Rand02] were compared with the results.
1.1 – GW estimation
The only dictated design parameter was a payload
weight of 440 lb (200 kg). Figure 1.1 presents GW vs. PL for
lightweight helicopters of which an initial estimation of
GW=1540 lb (700 kg) was derived. Helicopters with higher PL
and a much lower GW are not capable of two hours HOGE due
to smaller fuel capacities.
1.2 – RP estimation
Estimated GW yields a CP estimate of 170 hp (126.8 kW) (Figure 1.2).
1.3 – MDL and MRD estimation
500 1000 1500 2000 2500 3000 35000
100
200
300
400
500
N=22Plinear=0.107*Wo-0.435
PQuad=2e-5*W2+0.0321*W+55.5
Gross Weight[lb]
Pow
er[H
P]
Power vs. Gross Weight
Hellicopter DataLinear FittingQuadratic Fitting
Figure 1.2: Power versus GW
A large MRD (low MDL) means better hovering and
climb performance. Low MRD leads to low total weight, tail boom
length, balance, cost and ease of storage. From trend analysis an
initial estimation of MDL=2.9 lb/ft2 (138.9 Pa), subsequently a
MRD of 25 ft (7.6 m) was found.
1.4 – TRD Estimation
TRD should be large in order to minimize the TR RP.
Conversely, it should be kept small due to ground clearance,
minimum weight, and to prevent aft CG problems. From trend analysis, an initial estimation of TRD was 3.8
ft (1.2 m).
1.5 – Angular Velocities Estimation
As indicated in [Rand02], tip speed should be as high as possible for low rotor and drive system
weight. However, low tip speeds contribute to reduce noises. Correlation from [Rand02], yields an initial
estimate of Ω=470 RPM and ΩT=3150 RPM.
5
Section 2 – Airframe Design 2.1 – Design Criteria
The request for proposal (RFP) states that the airframe must be inexpensive to acquire. Since the
RFP requires the design of a trainer helicopter, safety is extremely important. Helicopter pilot trainees are
more susceptible to accidents and crashes than veteran pilots. Designing for crashworthiness minimizes or
even eliminates injuries and fatalities of occupants, as well as possibly alleviates some damage to an
aircraft’s structure that would normally be caused during a crash. Since crashworthy features contribute little
to the final cost and weight of an aircraft, crashworthiness design does not substantially alter the primary
criterion of inexpensiveness and is adopted as a design criterion [USAR89].
2.2 – Material
Materials for the Grass Chopper were chosen strictly from the manufacturing point-of-view to
maximize cost savings. The choices made are generally based on minimizing production time, part count,
and complexity and therefore will be discussed in more detail in the manufacturing section of the report.
Note that the main structural members of the Grass Chopper are made of graphite-epoxy or glass-epoxy
composite materials unless otherwise noted in the following sections.
2.3 – Major Features
The main airframe design is divided into two main parts: a cabin containing seats, flight controls,
and flight instruments and an aft part containing the cargo, engine, transmissions, and two fuel tanks. The
outer shape is designed with aerodynamic considerations in order to reduce drag where possible. A trade off
was made between the light weight design of an open aft configuration and the lower drag, closed aft design.
The closed design reduces drag by producing lower form drag during forward flight and reduces down force
in hover with only a minor weight penalty. Moreover, vehicle esthetics improves significantly with the
closed aft design, making the final product more attractive to potential customers. The airframe is moderately
curved to maintain a lower cross-section gradient necessary to avoid flow detachment.
2.4 – Structural Floor Design
The structural floor is a key component of the airframe design. It provides attachment points for
several parts. The structural floor was designed to have five longitudinal I-beams that provide stiffness and
attachment points for the crashworthy seats. The middle two beams extend farther out to allow for the
attachment of the tail boom struts and the landing gear. Five transverse members provide added stiffness to
the floor and also provide a means to crush the foam subfloor in the event of a crash.
2.5 – Subfloor Design
A subfloor is an important helicopter component because it dissipates kinetic energy through
crushing during a crash, which improves the occupants’ chances of survival [Jack99]. Two subfloor designs
were considered for the trainer GrassChopper. The first design was an aluminum alloy subfloor using keel
6
beams and intersection elements riveted together [Bisa02]. The second design consisted of five sections
made of Rohacell 31-IG foam and is based on a design from the 55th American Helicopter Society Annual
Forum and Technology Display [Jack99]. Although both designs are effective, the foam subfloor design was
selected because it is twice as light, made of inexpensive material, and is easier to manufacture than the
aluminum design. It weighs approximately 7.3 lb (3.3 kg) and requires less manufacturing time by
eliminating the need for over 400 rivets. Also, testing of a 1/5 scale fuselage model showed the foam design
exhibited excellent energy absorbing capabilities [Jack99].
2.6 – Landing Gear Design
The landing gear provides the protection and support for the GrassChopper when landing and on the
ground. Several landing conditions can occur as described in Table 2.1.
Table 2.1: Landing Conditions [Ligh88]
Landing Type Acceleration Descent Rate Requirements Normal 0-3 g 10 ft/s or less Small landing loads
Hard 3-6 g 10-20 ft/s Fuselage must not contact the ground Crash ~10 g 20-42 ft/s Prevent injury/death
These conditions are considered when designing the landing gear so that failure does not occur and injury is
avoided. A skid landing gear design was selected over a wheeled design because of its simplicity, lower cost,
and compatibility with the mission. The design of the landing gear was based on both the Schweizer 300Cbi
and the MBB BO 105 helicopters. Two separate skids are used for a small weight reduction compared to a
single piece design. Pivot attachments are located near structural members in the airframe, so that the landing
loads would be distributed to the structure. Oleo dampers around the pivot points provide a smoother
transition during takeoff and absorb impact energy.
2.7 – Tail Boom
An extruded aluminum alloy, monocoque design is chosen over two alternative configurations, a
semi-monocoque and a truss-type design. In addition to increased part count, riveting and welding required
in the semi-monocoque and truss designs respectively increase manufacturing time and cost. The extruded
metal tail boom design saves manufacturing time by eliminating the need to rivet and minimizing the need to
weld. Additional benefits of using an extruded tail boom includes minimal machining, no need for special
tools or jigs for assembly, and a new die can be made for only hundreds of dollars if the size is not available
[Nort05]. The extruded body is thicker to resist buckling [Dona93] resulting in sacrificed weight savings in
exchange for manufacturing cost savings.
The MR and TR radii plus a half a foot of clearance determined the length of the tail boom. The final
design is 12.5 ft (3.8 m) long, has a 6.5 in (16.5 cm) diameter, 0.125 in (0.318 cm) thick tubular cross
section, and weighs approximately 33 lb (15 kg). Examining the maximum velocity throughout the entire
7
flight envelope, the maximum Reynolds number of the tail boom is around 86,000; therefore the tail boom
will encounter fully laminar flow and will not experience sudden changes in Cd during flight.
The tail boom is attached to the airframe with a cantilevered mount and two support struts attached
1/3 of the way up the tail boom. The purpose of the support struts is to contribute additional support and
provide a location for active tail buffet damping actuators (discussed in section 6.6).
2.8 – Aft Fuselage Layout
The aft fuselage of the Grass Chopper contains the engine/drive train system, fuel tanks, and the
cargo area. The design layout was iterated using four major design criteria; these being a low average CG
location, a crashworthy design layout for occupant safety, good drive system performance, and provide a
storage location for 44 lb (20 kg) of miscellaneous equipment. Figure 2.2 illustrates four main iterations of
the design.
(a) (b) (c) (d)
Figure 2.2: Drive System Layout Iteration
The preliminary design (Figure 2.2a) was based on trends identified from observations of current
turbine helicopters. Although turbine engines are commonly attached to the top of the airframe, the engine
and gearbox mounted above occupants creates a hazard during crash scenarios. This design was modified by
placing the engine inside the airframe (Figure 2.2b) where there was unused space. The layout created a
lower CG location and better crashworthy performance by keeping large mass items low, but required a
complex structure to support the engine. The design was to keep the engine horizontal to simplify the
structure while still keeping the engine low for favorable crashworthiness performance (Figure 2.2c). This,
however, left little room for a storage and lead to a need for a complicated gearbox and long drive shafts
leading to driveshaft fatigue. Lastly, the final design has the engine higher then the previous design and aft of
the occupants for shorter drive shafts while maintaining a favorable crashworthiness design at some expense
of raising the CG (Figure 2.2d).
2.9 – Major Drive System Supports
The two major drive system support structures used in the Grass Chopper are the static mast and the
engine deck. These designs result from the chosen drive system layout and crashworthiness criterion.
8
2.9.1 – Static Mast Supports
A static mast is used as a result of the low drive system layout. This optimizes the use of the
bulkheads by applying rotor loads directly to the top of the bulkheads, instead of inside the airframe at the
transmission. A horizontal bar helps alleviate any inward bending of the main lift frame (large bulkheads).
The static mast supports also allow the attachment of a damper system to reduce vibrations created at the
MR. Additional discussion on the static mast is included in the gearbox section.
2.9.2 – Engine Deck
Large massive items must be prevented from intrusion into other vital parts of the airframe during a
crash scenario [USAR89]. The engine deck serves this purpose by preventing impact of the engine and MR
gearbox into the fuel tank, storage location as well as the occupant area in the event of a crash landing up to
10g’s [Ligh88]. The engine deck is constructed of horizontal box beams supported by bulkheads and lower
structural supports. This simplistic design allows for engine and MR supports to be attached onto the engine
deck easily and serves as mounting locations for firewalls.
2.10 – Bulkheads
The Grass Chopper’s bulkheads shape the fuselage, transfers exterior loads to the structural floor and
maintains the airframe’s structural integrity during rollover. Three major bulkheads used are two lift frames
and a rear engine bulkhead. The bulkhead’s circular shape attains favorable crash resistance characteristics at
the expense of some useable interior space, and have box beam cross-sections in order to provide high
compressive and bending strength [USAR89]. The rear engine bulkhead protects the engine during rollover
and allows a firewall attachment location between the engine and the tail boom.
2.11 – Cabin Structure
The primary function of the cabin support structure is to provide occupant protection. The side supports also
provide a load path for landing loads transmitted by the front landing gear attachments. Rotor blade intrusion
into the cockpit is a common cause of fatality during a crash; therefore, overhead beams positioned above
occupants prevent injury due to rotor blade intrusion [Cras05].
2.12 – Firewalls
FAA regulations, on rotorcraft 7,000 lb (3182 kg) and less dictate that occupants, all parts crucial to
a controlled landing, and baggage compartments are protected in the case of a fire [Fede05]. The firewalls in
the Grass Chopper isolate the cockpit, fuel tank location, baggage compartment, main gearbox, and engine.
Material design of the firewalls is described in the manufacturing section.
9
Section 3 – Turboshaft Engine Design 3.1 – Turboshaft Engine: Introduction
The GrassChopper is installed with a turboshaft engine (Model PSU250) that operates on Jet-A
kerosene aviation fuel. The engine is capable of delivering a maximum of 250 shp (190 kW) uninstalled at
ISA SSL and 222 shp (165.5 kW) at ISA SSL + 20°C. There are features of the PSU250 which keep it low
cost through each phase of production, purchase, and operation. These features are discussed throughout and
outlined in section 3.11. The design is very similar to the SOLAR Titan T62 auxiliary power unit proven
durable on many military aircraft and for ground electric power generation [Benini03]. Key features include
an automotive style startup, high performance automotive fuel injection, a FADEC fuel control system, and
innovative structural design which allow easy engine installation into the airframe.
In order to develop a conceptual design for a small scale turboshaft engine, it was necessary to
research other existing engines of similar scale. Existing engines that were studied include the auxiliary
power unit (APU) on the V-22 Osprey, the Rolls-Royce 250 C20W, the Boeing 502-6 turboshaft, and
turboshaft engines manufactured by Innodyn Turbines. Innodyn is a company located in Phillipsburg, Pa,
USA, that develops small scaled turboshaft engines for fixed wing aircraft. Their engine design, internally, is
very similar to the Solar T62 Titan APU. However, they have developed a new full authority digital engine
control (FADEC) system, coupled to an innovative fuel delivery system which greatly reduces purchase cost.
The core to the design is the patented automotive style fuel system which is capable of sustaining a
continuous combustion. They are currently selling their engines for between $26,000 and $35,000 dollars,
depending on which model is chosen. Innodyn agreed to share helpful information with Penn State to aid
their engine design, and Penn State agreed to share useful findings with Innodyn regarding installation of a
similar engine on a rotorcraft.
Since the request for proposal specified a conceptual engine design, the engine design is based on
preliminary thermodynamic cycle analysis and does not discuss component design. Conservative adiabatic
efficiencies are assumed. A MATLAB program was written to determine engine performance based upon
equations used in cycle the analysis found in [Hill91] and is found in Section 3.10.
3.2 – PSU250 Operating and Performance Specifications
Table 3.1: PSU250 Operating Specifications Mass Airflow Rate 2.2 lb/sec (1.0 kg/sec) Compression Ratio 3.5:1
Maximum TIT 1880 °F (1027 °C) Operating RPM 60,000 Maximum EGT 1315 °F (713 °C)
Table 3.2: PSU250 Performance Specifications Uninstalled Installed (assuming 10% parasitic losses)
MP 250 hp (186.4 kW) 225 hp (167.8 kW) MCP 225 hp (167.8 kW) 202.5 hp (151.0 kW)
SFC at MCP (lb/hp-hr) 0.712 0.712
10
3.3 – PSU250 Section View
Planetary Assembly
Turbine Compressor
Output to MR gearbox
Overrunning clutch
Inlet guide vane
Alternator Starter Combustion Area
Figure 3.1: PSU250 Section View
Engine mounts
11
3.4 – PSU250 Exploded View
Figure 3.2: PSU250 Exploded View
3.5 - Intake
The inlet system consists of an intake manifold with dual plenums facing out each side. A particle
catching screen wrapped around the circumference of the inlet guide vane traps large debris. An integral
mass airflow sensor alerts the pilot of any restricted airflow and is displayed on the instrument panel as an
indicator light. An optional upgrade includes particle separators, electrically driven and incorporated midway
between the outmost portion of the intake manifold and the inner housing. The entire inlet system can be
installed after the engine is mounted into place. It is made out of 0.125 in (0.318 cm) wall 6061-T6
aluminum, weighs around 10.0 lb (4.54 kg), and has an adiabatic efficiency of 0.92.
3.6 - Compressor
The PSU250 has a single stage centrifugal compressor, axial inflow and radial outflow. Detailed
design of the impeller will produce a constant 28.6 ft3/sec (0.81 m3/sec) volume flow rate into the combustor
when rotating at 60,000 RPM. The compressor has a compression ratio of 3.5, which is moderately low for
advanced compressor design in the past decades. A higher compression ratio would increase overall fuel
efficiency, but this would also increase the scale and power output beyond what the mission requires. The
efficiency of the compressor is 0.72 [Benini03]. Because the compressor is similar to the Solar T62
compressor, the manufacturing process is well defined and relatively inexpensive, and the research and
development process costs less than experimenting with higher efficiency compressors.
The impeller is made out of a dual-alloy titanium. The airfoil section is made out of a high
temperature, creep-resistant alloy while the hub assembly is made out of a high-strength, fatigue resistant
alloy. The cost of manufacturing these impellers will be no more than the current impeller costs in the T62
[Gayda98]. The diffuser housing is structurally capable of supporting the weight of the engine. The top and
12
bottom engine mounts are bolted directly to a flange running the circumference of the housing. The burner is
bolted to one end of it, and the inlet guide vane is bolted to the opposite end. In this way, the casing serves a
structural member and is a very compact way to integrate the engine into the airframe.
3.7 - Burner
The PSU250 has a reverse flow, annular combustor with eight automotive fuel injectors evenly
spaced around its circumference. Detailed design of the burner will meet a few general design goals, such as
high atomization, high swirl, stable flame, and sufficient secondary cooling air. Innodyn has developed a fuel
injection system that produces very small fuel particles at a very high pulse rate. This is the key cost
reduction feature of the engine. Instead of a traditional fuel system which delivers a continuous fuel flow, the
PSU250 makes use of discrete injection capable of sustaining sufficient atomization and power in the burner.
The burner is assumed to have an adiabatic efficiency of 0.98.
3.8 - Turbine
The turbine is designed to extract as much enthalpy as possible before exhaust. Part of the energy it
generates is used to drive the compressor, and another part is used for useful shaft horsepower. The
remaining enthalpy is exhausted as hot gas. It is a radial inflow, axial exhausting turbine and spins at
60,000RPM. It has an efficiency of 0.78. The turbine pressure ratio is designed to be about 0.225 and its
corresponding temperature ratio is about 0.759. Knowing mass flow rate, fuel consumption, and shaft
horsepower of an operating Solar T62, turbine properties were acquired through moving backwards through
the program [Benini03]
3.9 - Exhaust
The GrassChopper exhausts straight out the back of the aircraft. This was originally a concern
because the MR downwash is going to downwash exhaust gasses on the tail boom that are around 1315 °F
(713 °C) at MP. However, the shroud spanning the tail boom that encloses the TR drive is lined with a heat
shield, and any portion of the tail boom that is within the region of this hot downwash will need to be
covered. There will be additional heat diffusion from the cooler downwash of the MR which will dissipate
the hot exhaust gasses to an appropriate temperature.
3.10 - Engine Performance Analysis – MATLAB Program
A MATLAB program was written to analyze the operating performance of the proposed
GrassChopper engine. The MATLAB program was used to generate fuel performance curves which were
used to analyze the over all mission performance of the aircraft. Variations of the program were used to study
the performance gains and losses when certain design features were modified. It runs through a range of
maximum cycle temperatures and computes discrete points which were later plotted in Microsoft Excel.
Figure 3.6 is a trend graph of power to weight ratio versus specific fuel consumption. Data from
other turboshafts was taken from [Leyes99] for the trend points. Moving up and to the left, power density of
the engines increase. The PSU250 is on the lower end of the power density curve, and also the mid to higher
range of fuel consumption. Figure 3.7 is a trend graph of specific fuel consumption versus maximum
continuous shaft horsepower. The PSU250 is on the lower end of the power curve and as a result, suffers
from lower efficiency but benefits from lower production and operating cost.
Figure 3.3 is a graph of
specific fuel consumption versus shaft
horsepower for the PSU250 operating
at ISA sea level. Figure 3.4 is a
similar graph but for ISA + 20°C. Fuel
consumption at higher atmospheric
temperatures does not change much
with respect to shaft horsepower, but
available power drops off at higher
altitudes. This is shown more clearly
in Figure 3.5. Available shaft
horsepower drops linearly with
altitude. At the GrassChopper’s mission of 6,000 ft (1829 m) altitude and ISA + 20°C, the MP available is
about 193.0 hp (143.9 kW). Assuming 10.0% power transmission and parasitic losses, the available power to
the rotors is 173.7 shp (129.5 kW).
13
Figure 3.3: PSU250 – SFC versus SHP and Altitude, ISA, Jet-A fuel
0.65
0.7
0.75
0.8
0.85
0.9
100 110 120 130 140 150 160 170 180 190 200 210 220
SHP/h
p-hr
)SF
C (l
b
6000 ft5000 ft4000 ft3000 ft2000 ft1000 ft0 ft
14
0.65
0.7
0.75
0.8
0.85
0.9
100 110 120 130 140 150 160 170 180 190 200 210 220
SHP
SFC
(lb/
hp-h
r)
0 ft1000 ft2000 ft3000 ft4000 ft5000 ft6000 ft
0.5
0.6
0.7
0.8
0.9
1
1.1
100 200 300 400 500 600 700 800 900 1000
SHP
SFC
(lb/
hp-h
r)
Increasing cost, power, and efficiency
PSU250
145.00
165.00
185.00
205.00
225.00
245.00
0 1000 2000 3000 4000 5000 6000
Altitude (ft)
SHP
Figure 3.5: Maximum SHP versus Altitude, 1300K max cycle temperature
Figure 3.7: SFC versus SHP
265.00
ISAISA + 20C
0.50
1.00
1.50
2.00
2.50
0.5 0.6 0.7 0.8 0.9 1 1.1
SFC (lb/hp-hr)
Pow
er to
wei
ght (
hp/lb
)
PSU250
Increasing cost
Figure 3.6: Power to weight versus SFC
Figure 3.4: PSU250 – SFC versus SHP and Altitude, ISA+20C, Jet-A fuel
15
3.11 – Engine Cost Discussion
Schweizer Aircraft Corporation claims the installed cost of the Rolls-Royce 250 C20W installed on
the Schweizer 333 turbine helicopter is around $180,000. Innodyn’s most powerful model (255 shp) costs
around $35,000 to install. The PSU250 is estimated at around $60,000 installed into the GrassChopper.
The RR 250 C20W is much more
expensive because of higher efficiency, proven
reputation, and a higher rated power. The PSU250
is less expensive because of innovative design
alterations and less operating efficiency (see
Figure 3.8). The purchase cost of the PSU250 is
low enough to compensate for the decreased fuel
efficiency. At normal operating power, the
PSU250 consumes 17 gal/hr (64.4 liter/hr) of fuel
while the RR 250 C20W consumes 14 gal/hr (53.0 liter/hr) of fuel. Estimating Jet-A kerosene to cost $4.00
per gallon, the PSU250 uses $12.00 more per hour in fuel than the RR 250 C20W. This correlates to 10,000
hours of flight time in the PSU250 to match the cost of the RR 250 C20W, which has an average component
overhaul life of 2,500 hours. Purchasing the PSU250 at a lower initial cost and operating at a poorer
efficiency will be cheaper than the RR 250 C20W. Table 3.3 is a summary of the cost related features of the
PSU250, along with their advantages and disadvantages.
Figure 3.8: Engine Cost Comparison
0
50,000
100,000
150,000
200,000
Engine Cost ($)
InnodynRR 250 C20WPSU250
Table 3.3: Summary of cost related features for the PSU250
Feature Advantage Disadvantage Single stage centrifugal compressor backed to a single stage centrifugal turbine
Low cost, high power density, lightweight, increases TBO
Poor efficiency
Turbine/compressor unit cantilevered into engine by attached planetary gearbox
Eliminates bearings and oil in engine
Engine tolerances are coupled to gearbox tolerances, gearbox failure could damage engine
Performance automotive fuel injection
Low cost, adjustable operating parameters
Lower fuel efficiency
Diffuser as a structural member Compacts engine structure, simplifies engine installation
Diffuser structure is more complex
3.12 – Operation and Summary
The FADEC system controls the operation of the PSU250 from startup to shut down. It is monitoring
EGT, TIT, MR RPM, and mass airflow into the compressor. It uses these parameters to adjust the fuel flow
as necessary to provide the power demanded by both collective and cyclic piloting input.
If the engine speed falls below 10% of its normal operating speed, the igniter is lit and the pilot is
warned of the abnormal operation with an instrument light. This is in the case of a malfunctioning startup
16
sequence, or if the flight controlling requires more power than the PSU250 can deliver. Overspeed and
overtemperature protection reduces fuel flow to return the engine to its normal operating conditions.
In summary, the PSU250 is an extremely cost competitive turbine engine and fits nicely into the
GrassChopper package. It has a very low part count which makes assembly and disassembly relatively easy.
There is only one main rotating assembly, and this makes the internal operation of the engine simple,
maintainable, and durable. The unit itself is cantilevered into the engine to eliminate bearings and oil in the
engine. Using single stage centrifugal turbomachinery definitely reduces the cost of production and
maintenance, but there is a fuel efficiency decrease due to tip losses associated with centrifugal
turbomachinery. The innovative fuel delivery system developed by Innodyn Turbines drives the cost of the
engine very low. The performance automotive fuel injection greatly reduces the cost from conventional fuel
delivery packages while maintaining acceptable operating performance in the engine.
Section 4 – Gearbox Design 4.1 – Planetary Gearbox
There is a single stage, 5:1 planetary gearbox integral with the Grass Chopper engine. First, it is used
to cantilever the rotating assembly into the engine. This eliminates the necessity for oil and bearings in the
engine itself. Oil in the planetary gearbox is shared with the MR gearbox. The gearbox will have previsions
to ensure all critical bearings are properly lubricated, and that an uninterrupted circulation of oil persists
throughout the full operating range of the GrassChopper (see section 6.2). The main shaft cantilevering the
rotating assembly is rigid enough to withstand any lateral forces it will be subjected to.
The second purpose of this integral gearbox is to provide a 12,000 RPM shaft output to the MR
gearbox. The single stage planetary gearbox consists of a sun gear, three planet gears, and a ring gear which
is integral to the gearbox housing. The sun gear is spinning at 60,000 RPM and the planet gears are
connected by a planet arm which rotates at 12,000 RPM. The sun gear has a radius of 1.0 in (25.4 mm), and
each of the planet gears have radii of 1.5 in (38.1 mm). The ring gear has a radius of 4.0 in (101.6 mm).
Although detailed design of the gears was omitted, this initial sizing creates a 5:1 rotational ratio. The design
of the gearbox includes a combination of thrust and roller bearings that withstand axial and radial loads in
both directions. All of the loads transmitted to the output shaft of the MR gearbox are reacted inside the
planetary box, and none are transmitted into the engine.
The alternator and starter are attached to the planetary gearbox. The starter is incorporated before the
overrunning sprague clutch. The engine is started by an automotive starter and automotive spark plug. The
starter must be capable of turning the turbine at about 2,200RPM [Innodyn05]. This corresponds to a MR
rotation of 18 RPM prior to startup. It will be mounted before the over running clutch so when it turns it is
rotating the entire power train assembly. The alternator is incorporated after the overrunning clutch so that
during an autorotation, the aircraft still has full electrical capability.
At the given dimensions, the planetary gear reduction is 6:1 providing a MR output of 500 RPM.
The final stage in the main gear box is the tail shaft output. The stage is made of a 90° spiral bevel
set. The central shaft spins at 3,000 RPM and the pinion has a diameter of 2.2 in (55.9 mm). The output gear
has a diameter of 2 in (50.8 mm) and rotates at 3,300 RPM.
The TR gearbox was intended to be a simple design to minimize cost and weight. Since this
component is attached to a long moment arm from the center of the helicopter, minimizing weight results in a
more central CG. Minimizing the complexity also reduces costs, which is central to this helicopter design.
4.3 – Tail Rotor Gearbox
Gearbox design is constrained to the input and output criteria. On the Grass Chopper, the engine
output is 12,000 RPM. The MR output is 500 RPM and the TR output is 3,300 RPM.
Gearbox design is constrained to the input and output criteria. On the Grass Chopper, the engine
output is 12,000 RPM. The MR output is 500 RPM and the TR output is 3,300 RPM.
The main gearbox for the Grass Chopper has three stages: Engine input, tail output and MR output.
A centerline drawing of the drive system is included in Figure 4.2.
The main gearbox for the Grass Chopper has three stages: Engine input, tail output and MR output.
A centerline drawing of the drive system is included in Figure 4.2.
4.2 – Main Rotor Gearbox Main Rotor Gearbox
Above the input gear stage is the planetary gear stage. The stage operates with a locked ring gear and
input from the sun gear attached to the central shaft. The sun gear has a diameter of 1.5 in (38.1 mm) and 30
teeth. There are 4 planet gears, each with a diameter of 3 in (76.2 mm) and 60 teeth. The outer ring gear has
an inner diameter of 7.5 in (190.5 mm). The gear ratio of a planetary system with a locked ring gear is:
Above the input gear stage is the planetary gear stage. The stage operates with a locked ring gear and
input from the sun gear attached to the central shaft. The sun gear has a diameter of 1.5 in (38.1 mm) and 30
teeth. There are 4 planet gears, each with a diameter of 3 in (76.2 mm) and 60 teeth. The outer ring gear has
an inner diameter of 7.5 in (190.5 mm). The gear ratio of a planetary system with a locked ring gear is:
The engine input gear reduction step is comprised of a 4:1 spiral bevel at 90°. The input pinion,
coming from the engine, has 30 teeth. The input gear, connected to the central gear shaft, has 120 teeth. This
step re-directs the rotation and spins the central shaft at 3,000 RPM.
The engine input gear reduction step is comprised of a 4:1 spiral bevel at 90°. The input pinion,
coming from the engine, has 30 teeth. The input gear, connected to the central gear shaft, has 120 teeth. This
step re-directs the rotation and spins the central shaft at 3,000 RPM.
A static mast is used to carry the rotor thrust and
moment loading using two sets of bearings. A pair of duplex
ball bearings at the top react rotor thrust while roller bearings
react the moments. A conceptual drawing of the static mast is
pictured in Figure 4.1.
4.4 – Static Mast
The gearbox consists of a 90°, spiral bevel redirection. Since the TR driveshaft rotates at the required
TR speed of 3,300 RPM, no change in shaft speed is required, allowing the TR gearbox to be simple and
small, reducing weight and cost. Each shaft is held in place by a thrust bearing.
17
SRRG += 1..
Figure 4.1: Static Mast
18
4.5 – Drive System Schematic
Figure 4.2: Drive System Schematic
19
Section 5 – Rotor Design 5.1 Main Rotor
10 20 30 40 5040
60
80
100
120
140
Rotor Diameter [ft]
P [HP]
Required Power for Hovering Vs. Main Rotor Diameter
Thrust=700kgf(1544lb)Tip Speed=197m/s(646ft/s)NACA0012 Profile
NB=2NB=3NB=4
The primary design criterion for the MR was to satisfy the requirement of hovering at 6000 ft in
ISA+20°C conditions. However, autorotation requirements were also accounted for. Forward flight and
maneuvering efficiency was of secondary concern.
5.1.1 Theoretical Point of View
5.1.1.1. Hovering
From classical momentum theory: Vi ~1/r, and for hovering,
the induced RP is T⋅Vi. It follows that MDL should be kept as low as
possible for efficient hover performance (Figure 5.1).
5.1.1.2. Autorotation Figure 5.1: Required Power for Hovering versus MRDThe MR moment of inertia should be high to store a high
amount of kinetic energy and have a successful landing. In addition, during autorotation the rotor is driven by
air inflow, no power is supplied by the engine itself. For example in the case of vertical autorotation:
T⋅Vd=T⋅Vi+(profile-drag power). Therefore, the autorotation rate of descent decreases with MDL.
5.1.2 Number of Blades
A two bladed rotor was chosen for fixed thrust and tip speed, lower Nb means smaller RP, requiring
larger MRD, which leads to lower MDL and improves autorotation/hover performance (Figure 5.1).
5.1.3 Main Rotor Chord
5 10 15 20 25 300
5
10
15
20
MRD[ft]
RPI
[%]
Relative Performance Improvemet vs. MRD
MRC and MRD determine the solidity. Lower angles of attack are needed for larger solidity. MRC is
0.6 ft (0.2 m) for effective angles of attack (also Cl and L/D) along the blade, which, in turn minimizes
parasite RP.
5.1.4 Main Rotor Diameter
Larger diameters means lower MDL, improving
autorotation and hover performance. On the other hand, large a
MRD requires longer tail booms, thus shifting the CG significantly
backwards. The Relative Performance Improvement (Figure 5.2) is
defined asrr
VV
i
I
11≈
∂∂ . For a MRD>25 ft (7.6 m) the RPI is only
about 3%, hence a MRD of 26 ft (7.9 m) was selected.
Figure 5.2: Relative Performance Improvement versus MRD
20
5.1.5 Airfoil Section
The NACA23012 airfoil was selected for the GrassChopper. Main considerations were good
autorotation performance, hover performance, and control characteristics. From [Gess99], the Cd/Cl3/2 ratio
of each blade-element should be kept as low as possible for a minimum rotor profile-drag, leading to lower
rates of descent and improving autorotation. Using “DesignFoil” software for various airfoils, the
NACA23012 airfoil was proven to have the lowest values (Figure 5.3).
Cd/Cl^(3/2) for some checked airfoils. Re=1.75e6
0
0.01
0.02
0.03
0.04
0.05
0.06
0.07
2 4 6 8 10 12 14Angle of Attack[deg]
Cd/CL
(3/2
NACA 8-H-12
NACA 63-015A
: NACA 63-210 )
NACA 63-212
NACA 11-H-09
NACA 2415
NACA 5-H-15
NACA 63A010
NACA 64-008A
NASA/LANGLEY
NACA0012
NACA 23012
Figure 9.4: The NACA23012 has one of the smallest values of Cd/Cl
(examined using Designfoil), showing that this airfoil will also
improve hover performance. Lastly, NACA230 series have easier
control performance because they have relatively low Cm,
minimizing periodic blade pitch and undesirable periodic stick
forces, vibrations and control-position gradients.
0 0.2 0.4 0.6 0.8 10
2
4
6
8
10AOA distribution along the blade
r/R
AO
A[d
eg]
12[deg] washout NACA23012
Figure 5.5: AOA Distribution along Main Rotor Blade
5.1.6 Blade Twist
0 5 10 15 2094
95
96
97
98
99Rotor Required Power vs θ1
tw
θ1tw[deg]
Rot
or P
ower
[HP]
Figure 5.3: Cd/Cl^(3/2) for Checked Airfoils
Figure 5.4: Required Power versus 1twθ
Twist effect on hover performance was studied using
blade-element theory, including tip losses and compressibility
effects using Prandtl-Glauert correction. High blade twist means
better hovering performance (more uniform inflow distribution).
Optimal “1/ ” twist is very difficult to manufacture so the
cheapest, most common solution of linear twist was naturally
selected. High negative twist means large effective angles of
attack at the root leading to stall, therefore increasing RP while
reducing efficiency. A moderate, conventional -12° twist is used
(hovering results - Figures 5.4 and 5.5).
r
21
5.1.7 Taper
Blade taper and blade twist have quite similar to effects, both resulting in a more uniform inflow
distribution. Tapering also results in a several percent increase in rotor hover performance (lower RP),
however, this is not justified by additional production cost for a lightweight, cheap helicopter. Therefore, no
taper was applied to the GrassChopper.
5.1.8 RPM
Critical Mach numbers at the tip should be
prevented using tip sweep and supercritical airfoils. A
rather simple and cheap solution is limiting the tip speed
to 689 ft/s (210 m/s) ( ), thus not allowing the tips
to reach critical Mach numbers. Therefore, 500 RPM
was selected for a MRD of 26 ft (7.9 m), which is 30
RPM higher than the initial estimate (section 1.5).
0.6M ≈
y
x
Figure 5.6: Modeled Rotor Blade
5.1.9 Blade Structural Design
The GrassChopper’s blades are
all-extrude Al-2024 blades (section 5.9.1)
with: Sy=27 ksi (185 Mpa), E=9,850 ksi
(68 Gpa). The structure was modeled
using the two main I-sections (Figure 5.6)
considering both gravitational and
centrifugal forces.
Figure 5.7: Final Rotor Blade Design
On the ground, gravitational forces bend the blade leading to considerable stresses at the root. Root stress and
bending moment are given by: (1) ; (2)2
bb b
b
M y mgRMI
σ⋅
= = . In addition, tip displacement, given by
4mgR(3) =8EI
δ , ([Popo98]) is limited to 1.7 ft (0.52 m) according to FAA ground clearance requirements of
7.08 ft (2.16 m) ([AMCP74]). Centrifugal forces cause tension stresses along the blade with a maximum at
the blade root given by:21(4)
2tb b
CF M RA A
σ Ω= = . Finally, for the TR design, δ=0.76 ft (0.23 m); σb=13ksi
(89,630 kPa) σt=4.1ksi (28,270 kPa).
22
5.2 Hub Design
MR has two blades connected to the hub as a seesaw allowing for blade flapping. Overall hub
simplicity and reduced bending stresses was of primary importance, also allowing us to use smaller parts.
Trailing edge flap system control was considered, but was found inappropriate due to complexity and cost.
Therefore a rather conventional swash plate was designed, minimizing costs while maximizing reliability.
5.3 Anti-Torque System
5.3.1 The Choice of Anti-Torque System
The above considerations also lead to a TR design, chosen as the GrassChopper's anti-torque system.
The other two alternatives were: a small side wing connected to the tail boom (using the MR downwash) and
a laterally ejected jet (from the end of the tail boom). The first was rejected due to complicated aerodynamic
analysis. The second creates a lateral force using air jet momentum, thus eliminating a TR, driveshaft, and a
gearbox, but requiring additional construction (increased GW), and an air compressor to provide the jet mass
flux (again, significant GW increase). Bleeding the engine’s gas generator or using the engine exhaust are
possible; however, bleeding is limited by engine required specific pressure ratios and the engine’s exhaust
for this case may not be strong enough. In addition, the complexity of the overall system is significantly
higher. Lastly, electrically and hydraulically driven tail rotors were researched but found to be too heavy for
this application. The above mentioned considerations led to a conventional TR.
5.3.2 Tail Rotor Design
The TR was designed as a pusher rotor with similar design parameters as the MR. Aerodynamic
considerations were of less importance due to a RP of about 5-7% of the total RP. Manufacturing simplicity
and low price were the main parameters. Twist results in some performance increase but due to an already
low RP, this was not justified by additional cost, so no twist was applied. The final design is a two bladed
NACA 0012 rotor, TRD of 4 ft (1.2 m), TRC of 4 in (10.2 cm), ΩT=3,300RPM, and a tip velocity of 689 ft/s
(210 m/s) ( 0.6M≈ ).
Section 6 – Systems 6.1 – Fuel System Design
A major cost reduction design concept of the engine is an innovative fuel delivery system. The
engine runs on Jet-A kerosene fuel. There is a Full Authority Digital Engine Control (FADEC) unit which
manages the operation of the engine from startup to shut down. The fuel system is modeled off of an
automotive fuel injection system. The injection manifold consists of seven electronically controlled fuel
injection nozzles attached to the fuel injection manifold ring. High quality automotive fuel injectors will
provide a droplet size small enough for proper combustion in the burner, as well as a pulse rate high enough
to simulate a continuous fuel flow seen in existing turboshafts.
23
The engine is equipped with two high pressure electric pumps, each capable of providing at least 65
psi (448 kPa) of fuel pressure at the required fuel flow of 3.0 lb/min (1.37 kg/min) at MP. Only one pump is
used at a time to provide fuel to the injection nozzle manifold ring. The other pump is used as a backup and
is switched to automatically if a pressure drop is detected by the computer system. The pilot is noted of this
switch on the instrument panel. A high pressure regulator is used to maintain pressure in the manifold at 65
psi, and another pump is required inside the fuel tank and this must supply the high pressure pumps with at
least 2 psi (13.8 kPa) of fuel pressure when the flow is at its maximum rate. This is required because the fuel
tank is below the engine and gravity cannot be used as a fuel feed [Innodyn05].
The FADEC system will monitor fuel flow, fuel pressure, collective pitch, exhaust gas temperature,
and MR RPM. These variables will be used to control the fuel flow, and maintain a constant turbine RPM
throughout the power range required during flight.
6.2 – Oil System Design
The basic components of the GrassChopper oil system
are the scavenge pump, sump, main filter, magnetic particle trap
(MPT), cooler, fan, and main pump shown in Figure . The MPT
is located right after the main oil filter, and there is another filter
at the MPT output. Both filters have integral alarms and any
abnormal operations will be alerted to the pilot. When enough
chips are present to complete the electrical circuit, a warning
light will appear to the pilot. The main pump is electrically
driven and submersed in the sump, driving constant volume into
the filter.
Main GB Planetary GB
Scavenge Pump
M P T
Oil Cooler
Sump
Main Pump Filter
Figure 6.1: Oil System Schematic
Preliminary thermodynamic calculations were used to estimate the oil capacity and circulation
specifications necessary for the proper and safe cooling of the oil system. The two gearboxes which use
circulating oil are the MR gearbox and the engine-attached planetary gearbox. Assuming a worst case
efficiency of these two gearboxes to be 97%, and a 100% power condition of 225 hp (167.6 kW), this
requires 4.50 hp (3.35 kW) to be dissipated as heat. Assuming that the combination of the gearbox housings
and the sump will dissipate 10% of this heat generation through convection, this leaves 4.05 hp (3.02 kW) to
be dissipated by the oil system. The mass flow rate of oil through the system is given by
)( outinp TTcqm
−⋅=& .
Assuming use of MIL-L-23699 standard oil with a specific heat of .455 BTU/lb-°F (1904.2
J/kg-K), a density of 7.8 lb/gal (1.1 kg/L), and a temperature drop of 100°F (311 K) through the cooler, the
GrassChopper needs to circulate 0.48 gal/min (1.81 L/min) through the system. FAR §27.1011 requires that
24
there be at least one gallon of oil for every 40 gal (131.2 L) of fuel capacity in a certified rotorcraft. The
GrassChopper will have 2 gal (6.6 L) of circulating oil at about a four minute recirculation rate. If there is an
oil leak at the circulation rate, the pilot will have four minutes to land the aircraft before dry run which is
damaging to the gearboxes.
6.3 – Cockpit Control Panel
The control panel's shape is designed to enable maximum view range for the pilots while
maintaining indicators at appropriate heights and distances from the pilots. Indicators are shaded to prevent
reflections from the panel. The main references for the GrassChopper control panel is the Bell 205.
Considering missions, cost, turbine engine, and weight requirements, the GrassChopper will be equipped
with the following:
• Six warning lights: clutch, MR temp, MR chip, TR chip, starter on, low fuel.
• Flight instruments: Altimeter, Ball, Magnetic Compass, Vertical Speed, Tachometer, VOR, Air Speed
and Attitude indicators, all 3 1/8'' size.
• Engine gauges: Torque, RPM, EGT all 2.25in (5.7cm) diameter. Indicators include fuel pump, fuel filter,
ignition, engine oil temperature/pressure, gearbox oil temperature, and a voltmeter.
• Communication system: one radio and internal communication control panel.
6.4 – Crashworthy Seats
Crashworthy seats are vital to the safety of the pilot and student for this trainer helicopter. FAR part
27, guidelines for normal category rotorcraft, outlines specific guidelines for crashworthy seats. These seats
are to be designed so that persons using the seat will not suffer adverse consequences in the case of an
emergency. These seats must also have a safety belt and a shoulder harness with a single point release and
must also allow the pilot to exert full control of the controls. The regulations also state that the seat should be
able to support a person of at least 170 lb (77.3 kg). However, crashworthy seats should be able to support
people of a wide variety of weights and heights. Martin-Baker has designed a lightweight, crashworthy seat
that meets these requirements. This seat, the S-92 Crew Seat, accommodates 5-95% of both male and female
crew. It is adjustable in horizontal and vertical directions to accommodate for height differences. The cost of
each seat is approximately $200 [Mart04].
6.5 – Flight Control System
The flight control system was designed so that the linkages go aft between the two seats then
upwards to the hub. As in most helicopters, but especially for trainers, there are two sets of controls, one for
instructor and other for student. Conventional mechanical flight controls were used for cost savings.
25
6.6 – Active Tail Buffet Damping
Two electronic actuators acting as active control dampers are attached in line with the tail boom
support struts. Using an electronic control system and accelerometers mounted onto the tip of the tail boom,
tail boom buffet is monitored and reduced by constant feedback response from the actuators. Since tail
buffeting is created by white noise oscillations, the actuators require a large bandwidth to react to a wide
range of oscillation frequencies. Electric actuators are chosen over piezoelectric actuators to save cost and
over a passive offset mass system to keep a forward CG location. Furthermore, the electronic actuators can
be governed electronically to simulate elevated tail boom buffeting for training purposes [Alkh03].
6.7 – Landing Gear Systems
There are two simple additions to the landing gear which have been noted as extremely important by
rotorcraft instructors. The first important addition is to have wheels on the landing gear. These wheels have
two positions. To lock the wheel’s down, one person can push down on the empennage near the tail rotor
which raises the front part of the landing gear and the wheels can be lowered. This must obviously be done
when the GrassChopper is not in operation. Having wheels allows the helicopter to be moved around the
airfield and into or out of hangers while not in operation. A second addition is to have caps which can fit
onto the fore or aft ends of the landing gear. This helps maintain control over the CG location depending on
whether or not there is a student in the aircraft or the instructor is flying alone.
6.8 – Available Upgrades
One system that would be available as an upgrade is a variable stability & control system. The
United States Naval Test Pilot School has installed such a system in an SH-60B helicopter for in-flight
simulation. This aircraft, like the GrassChopper, uses mechanical flight controls and is capable of varying
stability & control parameters about the pitch, roll, and yaw axes [Mill94]. This type of system could be very
valuable as a training tool because several different flying conditions could be simulated without putting the
pilots or helicopter in danger. The ability to simulate non-ideal conditions is just as important as training
pilots to deal with normal flying scenarios.
Another available upgrade to the GrassChopper is a cockpit airbag system. Trainer helicopters are
more susceptible to crashes than other helicopters due to the nature of their mission. Preventing injury and
protecting life are the primary concerns when considering a crash scenario. More than 50% of fatalities in
helicopter accidents are caused by head strikes, especially when the pilot’s head impacts the cyclic control
[Cock06]. In order to increase safety and prevent injury/death, Armor Holdings Aerospace & Defense Group
has developed a cockpit air bag system (CABS) that consists of two forward and two lateral air bag modules
and an electronic crash sensor unit. The system can be modified to fit any aircraft and weighs only 23 lb
(10.5 kg) [Cock06]. The system has a predicted 0.0004 maintenance hour to flight hour ratio, a 60 second
data recording capability, and a WindowsTM operator interface [Cock06]. This system would be a vital asset
26
to any helicopter, but especially a trainer. Increasing the chances of survival during a crash is certainly a
worthwhile investment.
Section 7 – Performance Analysis 7.1 – Download-Force Estimation
The download created from rotor-fuselage interaction was estimated using the method found in chapter 4 of
[Prou95]. The method adds up drag fractions for each airframe segment, considering dynamic pressure ratios.
Download force to the GW ratio is calculated from:
( )1
2 . .. .
n
N
D nnV n
qC AD LDG W A
==∑
. The result for the case of a cylinder similar in size: ..
80WG
DNtD V
V ⇒= =
1.2% was validated against the theoretical estimation: dtheory SDvD 2
21 ρ= = 76Nt (a minor 4Nt deviation).
7.2 Trim Analysis
Trim analysis included solving the GrassChopper’s equilibrium equations for hovering and forward
flight velocities. The equations can be divided into the following four sets:
1. Three forces equilibrium equations.
2. Three moments equilibrium equations.
3. Three equations describing the flapping angels of the MR blades.
4. Glauert's equation, describing the inflow ratio through the MR disc.
The equations were used under the following assumptions:
1. MR trust equals the helicopter’s weight plus down force drag.
2. Small angles assumption ( xxx == ~)sin(,1~)cos( ).
3. Airframe drag was estimated using its cross – section and Cd=0.3.
4. only longitudinal flight was simulated (no side-slip or lateral flight).
A main interest was to find the following parameters as a function of forward flight velocity: three pilot
command angles, three MR flapping angles, two body angles, TR force, and the inflow ratio. These results
were used to determine other parameters in the design.
The equations are coupled and non linear, hence no analytic solution is available for them except for the
case of hovering. The equations are solved numerically using a specially designed multidimensional
unconstrained nonlinear minimization solver (employing the Nelder-Mead method). An initial guess for
forward flight numerical solution is the analytical solution for the case of hovering. Each solution step uses
the previous output as an initial guess.
After consulting a few helicopter pilots, longitudinal and lateral body shouldn't exceed 5°. Hence the
horizontal stabilizer was designed to provide a positive pitch moment and partially overcome rolling moment
27
produced by the TR. Its location is on the tail’s left hand side (same as the TR), 12 ft (3.6 m) behind the MR
shaft. It has a 2° negative incidence angle and a NACA 0012 airfoil, mainly due to simplicity and
convenience. It has a 2.5 ft (0.76 m) span and a 2 ft (0.61 m) chord. Note that the longitudinal pitch angle is
defined positive when the GrassChopper has a “nose down” pitch.
It is important to add that a vertical fin was not a part of the simulation, although all equilibrium
requirements were still satisfied. However, in the case of a TR shaft failure a vertical fin will provide the
necessary anti-torque at a certain forward speed which will allow the pilot to maintain control of the
GrassChopper, increasing the helicopter crashworthiness. This is a suggested addition to the design.
7.3 Performance Table 7.1: RFP Mission PhasesThe main RFP requirement is HOGE at 6000 ft for two
hours. Trim analysis results at that altitude were used to estimate
the required power. Using a fuel consumption of 0.8 lb/(hp-hr),
an estimate for the total required fuel weight was determined.
Table 7.1 divides this main RFP mission into five phases. In
conclusion, the GrassChopper requires a total of 276.3lb of fuel.
Mission Phase Time (min) Fuel lbsWarming up 2 13.1 Climb 6.1 13.1 Hovering 120 194 Descent 3.05 3.41 Reserve 20 52.5 Total 151.2 276.3
0 20 40 60 80 1000
50
100
150
200
250Power In Forward Flight
Speed [knots]
Pow
er [H
P]
Shaft PowerInduced PowerBody Drag PowerBlade Profile PowerEngine Power (losses=30%)
Maximum Indurance230 min at 60knots
Maximum Range290 miles at 89knots
Engine power analysis takes
into account induced power, airfoil
parasite drag, airframe aerodynamic
drag, assuming 10% installation power
losses. Figure 7.1 presents this analysis
results in a power vs. forward flight
velocity curves. The purple curve
represents the total output power
required from the engine at each
velocity. Out of this figure, the
GrassChopper’s performance figures
were calculated which were compared
to other helicopters in the same
category (Table 7.2 for an example). Figure 7.1: Power in Forward Flight
Helicopter Max. Velocity* Max. Range Max. Endurance
GrassChopper 115 [knots] 290 nm @ 89 knots 230 min @ 60 knots Robinson R22 103 [knots] 250 nm @ 90 knots 210 min
Table 7.2: GrassChopper Performance compared the to Robinson R22
*Calculated according to the simulation without considering maximum transmission rating.
28
Maximum rotor pitch is mechanically limited by 24° (Figure 7.2), hence maximum velocity is 115
knots. A simple turbine engine model for temperature changes with altitude was used. Figure 7.3 features
payload versus range. Note that with a maximum payload of 440 lb
(200 kg), the GrassChopper can reach up to about 290 nm. For
smaller payloads the range may increase up to about 320 nm for a
light-weight pilot of 150 lb (68.1 kg). Figure 7.3 was derived
assuming a fully fueled helicopter, and since the fuel weight
fraction (almost 20%) is quite large, the change in maximum range
is quite small (since the total change in GW is relatively small).
290 300 310 320 330 340 350
50
100
150
200
250
300
350
400
450
500
Range [nm]
Payl
oad
[lb
110 112 114 116 118 120 120
2
4
6
8
10
12
14
16
18
20
Max. Velocity [kt]
Alt
[Kft
]
Max. Velocity vs. Alt (ISA+20) Mechanical Limit - 115 [kt]
-40 -30 -20 -10 0 10 20 30 40 500
2
4
6
8
10
12
14
16
18
20ISA ISA+20 ISA+351300
1400
1500
16001700
W [lb]
OAT [Co]
Alt
[kft
]
HOGE vs altitude
]
Primary mission capability verification can be seen in
Figure 7.4 presenting HOGE altitude vs. OAT for different weight
configurations. Note that weight range exceeds GrassChopper
minimum and maximum values for the sake of showing trends.
The GrassChopper has a relatively high hover altitude ceiling,
but that is related to a slightly over-designed engine. For
example, hover altitude for 1500 lb (682 kg), on an ISA+20o
day, is about 12,000 ft (3,658 m), which is well over the basic
RFP requirement of 6000 ft hovering altitude.
Payload vs. Range (ISA+20o)
Maximum Payload
Figure 7.5 combines maximum velocity limitation (115
kts) along with the available altitude for each speed at the higher
end of the GrassChopper speed range. As speed is gradually
increased from 110 to 115 kts, the maximum altitude drops from
20,000 ft (6,096 m) to about 15,000 ft (4,572 m), until the performance curve meets the red line indication
maximum velocity.
Figure 7.2: Command Angles
Figure 7.3: Payload versus Range
Figure 7.4: HOGE versus altitude Figure 7.5: Max. Velocity versus Altitude
29
Section 8 – Manufacturing 8.1 – Manufacturing Techniques
8.1.1 – Electron-Beam Curing
Electron-beam curing is the proposed process to manufacture the composite materials in the
GrassChopper. This method of curing is a type of radiation curing that uses high speed electrons that collide
with atoms in a polymer-initiator mix. The high energy of the electron beams and the x-rays generated from
these beams penetrates into the composite, giving a uniform cure to materials [Lopa99].
Electron-beam curing is chosen over the traditional use of the autoclave and resin transfer molding
because it minimizes the temperature required to form composites, and it greatly reduces the use of volatile
materials required for curing. The high temperatures required for the autoclave and resin transfer molding
techniques can produce residual thermal stresses in the composite, and require resins that are not as stable at
room temperature as the resin required for the electron-beam curing. Furthermore, electron-beam curing
minimizes cost of handling, storing, and disposing of the materials [Lopa99].
The most valuable benefits of using Electron-beam curing come from the capability to manufacture
large, near shape components. This eliminates the use of expensive tools required to make smaller, precision
composite parts and minimizes the need for mechanical fasteners and adhesive bonds. Since manufacturing
large parts reduces the weight, part count, manufacturing time and manufacturing complexity, it significantly
cuts manufacturing costs. In fact, in the RWSTD program, Sikorsky Aircraft and the U.S. Army Aviation
Applied Technical Directorate proved that electron beam curing leads to a manufacturing cost savings
between 25%-50%. In addition to the cost savings, Sikorsky demonstrated a 60% reduction in tooling costs.
Lastly, composite repairs could be performed quicker and easier than traditional methods if small electron-
beam curing devices would be used onsite with the helicopter needing a repair [Lopa99].
8.1.2 – Paint-less Finish
A paint-less technique involves laying up parts with a pigmented surfacing film on the exterior
surface. This is another concept demonstrated by Sikorsky Aircraft in the RWSTD program. When the part is
cured, it has a colored finish that eliminates the need for primers or paint. Airframe weight can be reduced by
approximately 2% compared to priming and painting parts. This can be directly related to significant savings
in manufacturing cost and time [Kay02].
8.1.3 – Tool-less Assembly
Many parts and sub assemblies come from different companies and countries and need to be
produced so they are compatible. This has created the practice of using a large number of dedicated
manufacturing jigs and assembly tools which is undesirable. Cost can be lowered by tooling features that are
incorporated into the geometry of the component parts. Part variability results in gap or interference
conditions at the assembly level, otherwise the parts do not fit correctly. These parts are designed with loose,
30
adjustable features, both of which increase part count, weight and manufacturing labor. Using tool-less
assembly can lower costs for manufacturing, weight, and manufacturing time savings associated with
elimination of mechanical fasteners by utilizing detailed part features to locate parts within their assemblies.
This concept was proven effective in the RWSTD program (Sources from: [Burley98] and [Sandy03]).
8.2 – Component Materials and Concepts
8.2.1 – Rotor Blades Materials and Manufacturing
Materials commonly used for blades are: composite, aluminum alloys, and titanium. From these
choices, the main and TR blades are made out of aluminum alloy in order to maintain a low price. The price
range for aluminum blades, shown in Table 8.1, is due to different manufacturing processes: $900 for riveted
blades and $1,500 for extruded blades.
The all-extrude blade was chosen because of its lower drag (compared to the riveted blade which
does not have the smooth surface as the all-extrude blade), higher strength qualities, and because rivets are
known to cause fatigue cracks resulting in blade failures. The cost difference of $600 is worth the advantages
and it is a relatively low price difference compared to the other two options discussed in Table 8.1.
Table 8.1: Rotor Blade Materials Comparison
Advantages Disadvantages
Composites High strength Specific stiffness to weight ratio Bending-torsion coupling.
High cost (~$5,000), complicated manufacturing and repair processes, water absorption, and low inertia.
Aluminum Stiffness, high inertia, lower cost (~$900-$1,500), simple repair and maintenance.
Corrosion problems
Titanium Similar to aluminum Higher cost (compared to aluminum)
8.2.2 Airframe Materials and Manufacturing
Composite materials for the airframe are substituted extensively for traditional metals. Relative
advantages and disadvantages are given in Table 8.2.
Table 8.2: Airframe Materials Comparison
Advantages Disadvantages Composites High strength, specific stiffness, saves up to 30%
weight, reduced manufacturing cost, simpler assembly process, corrosion resistant
Damaged parts usually needs replacing (no simple repair and maintenance), conservative market, high initial cost (not an issue).
Metals Inexpensive and well known manufacturing repair & maintenance processes
Relatively high cost for material, higher weight, requires fasteners (added complexity)
The airframe skin material are made of fiberglass because of its relatively low density, and low cost
(compared the commonly used Graphite/Epoxy). Furthermore the skin does not require high strain range
therefore it is natural to choose fiberglass with a simple wet lay up process.
31
The GrassChopper’s windows are made of polycarbonate which is optically clear, providing
excellent total luminous transmittance and a very low haze factor. Being tough and lightweight it is ideal for
"see-through" applications where impact resistance is important, and another advantage is that it maintains its
properties over a wide range of temperatures from -40° F to 280°F [KMac06].
8.2.3 – Composite Airframe
Although aluminum has been preferred in the past because of its superior mechanical properties, the
forgings have lead times and tooling costs that exceed schedules and budgets of many programs. There are 2
forms of forging, each having drawbacks. The open method repeatedly manipulates parts with unconfined
rollers or dies until the desired method is achieved. This method is very inexact and time consuming. The
second method, closed forging, confines the material within a die cavity and uses pressures to aid material
flow into the unfilled portions of the die. This method is more accurate than the first but is dependent on the
quality of the die. Normally several dies must be made before the desired product is formed correctly which
adds to manufacturing cost and time. Composites are also significantly lighter than aluminum which have
showed a reduction in weight of about 20% (Sources from: [Kay02], [Lopata99], and [Moore03]).
8.2.4 – Foam Sub-floor
Using a foam sub-floor has several advantages over a common aluminum sub-floor. The original
aluminum alloy 7079 sub-floor weighted 16.62 lbs. By using Rohacell 31IG foam the weight is reduced to
7.3 lbs, a weight reduction of 56%. The original design had many parts, and required least 500 rivets. Using
the foam eliminates most of these rivets making manufacturing faster and cheaper [Jack99].
8.2.5 – Integrated Ceramic Composite Firewall
The integrated ceramic composite firewall material and process incorporates ceramic fiber and film
adhesive into one unique material. The technology reduces cost, part count, weight, and manufacturing
complexity. Existing firewall technology utilizes individual metallic panels as shields for fire penetration
protection. This requires secondary bonding and mechanical fasteners to insure proper installation to the
composite substructure. The bond line between the metallic panels and the graphic structure experience dis-
bonding and allow fluids to migrate between the two surfaces, creating potential safety issues [Misc05].
8.2.6 – Advanced Composite Joint Concept
Conventional methods of bolting or bonding have distinct disadvantages. Bolted joints are heavy and
labor intensive. Bonded lap joints rely on one of the weakest aspects of composites, inter-laminar tension. By
using interlocking finger joints these disadvantages can be avoided. This concept is intended for joining thin
laminate skins. The “fingers” are formed by water jet cutters and joined to form a continuous sheet with high
specific joint strength compared to the other 2 methods listed above [Kay02].
32
8.2.7 – Extruded Aluminum Tail Boom
Semi-monocoque structures and welded truss systems are commonly used for tail booms in
helicopters. These structures are time consuming to produce. Both structures have high part counts, the semi-
monocoque system requires many fasteners for its numerous supports, and the welded truss system requires
extensive welding at odd angles that cannot be done by machines.
The concept of using an extruded aluminum tube eliminates the complexity of manufacturing.
Extruded aluminum tubes are commonly stocked items that are cheap. Even if the desired tubes are not
available, new dies can be made for only a couple hundred dollars [Nort05].
8.2.8 – Summary
Most manufacturing methods and materials chosen are based on new, emerging technologies tested
and validated through research and testing. Table 8.3 lists some of these newer technologies and their
benefits, quality enhancements, and appropriateness to be used in industry (on a scale from 1 to 5, 5 being
ready for implication or currently being used).
Table 8.3: Manufacturing Techniques and Materials Summary
Method
Less Labor Time
Lower Labor Cost
Lower Part
Count Less
Weight Readiness
(1-5) Quality (↑,-,↓)
E-beam Curing 5 ↑ Foam Sub-floor • • 5 - Tool-less Assembly 3 ↑ Ceramic Composite Firewall • 5 - Advanced Joints 4 ↑ Paint-less Finish • 4 -
Section 9 – Weights and Center of Gravity Location This section describes the methods used to estimate the empty and maximum gross weight of the
GrassChopper, and to calculate its CG location, which is important for determining the trim and performance
of a helicopter.
9.1 Weight Estimation
The weights of the major components were first determined using the method formulated in Chapter
10 of Prouty’s text [Prou95]. This method requires initial parameters derived from the initial trend analysis
data. Next, the weights were estimated using analysis, as in the fuel weight analysis detailed in section 7.3, as
well as searching the market for similar components and valuating the weights using a SolidWorks program.
A listing of weights is included in the MIL-STD-1374 statement in Appendix A.
33
9.2 C.G. Estimation
-4 -2 0 2 4 6 8 10 12 14 1-10
-8
-6
-4
-2
0
X axis [ft]
Z ax
is [f
t]
155
57.3 44
298
130
44.1
99
105
143.3
397enginepilots and chairscargofuel tankmain rotortail rotoravionicsLanding GearTail
A helicopter's trim
and performance is influenced
directly by its CG location.
CG movements can occur due
to various mission definitions,
different pilots, cargo, fuel
quantity, and fuel
consumption during flight. In
this analysis fuselage stations
and waterline origin were
referenced to the center of the rotor
hub as shown in Figure 9.1. The
fuel is located underneath the main hub, which maintains the helicopters stability characteristics while the
fuel runs out (based on chapter 10 of [Prou95]).
C.G = (0.934,0,-3.27) (ft)
Figure 9.1: CG Location
Section 10 – Cost Analysis 10.1 – Description and Validation of Cost Model
The cost model used for the design of the GrassChopper was the given in the 2002 RFP where the
design was for a light helicopter upgrade. The term “light helicopter” was given to older four to six place
turbine helicopters. It was therefore assumed that this model could be used to accurately predict the cost of a
two place turbine trainer helicopter.
The cost model consisted of thirteen subsections which are as follows: rotor system, airframe,
landing gear, powerplant structure, air induction, propulsion, flight controls, instruments, electrical, avionics,
furnishings and equipment, air conditioning, and final assembly. Each subsection had an equation which
depended on the weight of the particular part of the helicopter and for specific subgroups there were specific
material factors. This model was highly dependent on the weight of specific components of the
GrassChopper. Therefore it is necessary to reduce the weight in order to reduce the cost.
This cost model was used for a four to six place turbine helicopter. While this is a small helicopter, it
was necessary to apply the model to other trainers for accuracy. The cost model was therefore applied to the
Robinson R22 and the Schweizer 300C, which are both piston engine trainers. The dimensions for each of
these helicopters were found in the appendices of Prouty’s Helicopter Performance, Stability and Control
(2002), where there was also weight breakdown which utilized the dimensions and determined a theoretical
weight. Once the specific weights for each helicopter were determined, they were utilized in the cost model.
34
The current cost of a R22 is $210K and the cost of a Schweizer 300C is $215.8K. Based on these
values, there was a 0.00% error for the Schweizer 300C and a 1.30% error for the Robinson R22. These
errors were significantly low and therefore it was determined that this cost model would accurately predict
the cost of the GrassChopper. There was some concern, however, that the cost prediction was “too accurate”.
It was therefore noted that the cost model may have been created from Prouty’s Text or these models.
10.2 – Recurring Cost Breakdown Table 10.1: Recurring Cost Breakdown The thirteen cost model subsections took into
account the factors which are associated with
recurring costs. The factors of recurring costs are
production labor, direct materials, process costs,
overhead and outside processing. Because there was a
significant focus on manufacturing methods, the cost
model was applied to two designs. The first cost
estimate was a model that uses all metallic materials
and the second model incorporated composites
heavily. Table 10.1 shows the recurring costs for the
aluminum and composite models.
Subsection/Cost Composite Metal Rotor System $ 12,400 $ 12,400 Airframe $ 59,700 $ 92,800 Landing Gear $ 11,200 $ 11,200 Propulsion $ 70,000 $ 70,100 Flight Controls $ 1,100 $ 1,100 Instruments $ 3,900 $ 3,900 Electrical $ 700 $ 700 Avionics $ 18,900 $ 18,900 Furnishings/Equipment $ 900 $ 900 Air Conditioning $ 3,400 $ 3,400 Final Assembly $ 44,600 $ 50,000 Total Cost $260,500 $275,000
It can be seen from Table 10.1 that the
composite model is approximately $15,000 less than the
aluminum version due to significant reductions in the
airframe and final assembly. There was a $30,000
reduction in the cost of the airframe and a $5,000
reduction in the final assembly of the GrassChopper.
Figure 10.1 shows a pie chart of the cost breakdown for
the composite model.
Rotor System, $12,400
Airframe, $59,700
Final Assembly, $44,600
Propulsion, $70,000
Instruments, $3,900
Avionics, $18,900
Flight Controls, $1,100
Landing Gear, $11,200
Furnishings and Equipment,
$900
Air Conditioning,
$3,400
Electrical, $700
10.3 – Cost Record
Piston helicopters are significantly less expensive their
turbine equivalents. The most common piston trainer
helicopters were approximately $210K. The hope was to design a helicopter which would not exceed $400K,
which was highly dependant on the type of engine selected. The first few months of the design process took
into account the costs well-known industry engines. The most significant reduction in cost, as seen in Figure
10.3, was the innovative design for the PSU-250. At that point in time the weights were theoretical estimates.
Another drop in cost in February was due to the reduction of MR blades from four to two. As the detailed
design began to develop, most accurate and higher weight estimates were determined and the cost began to
Figure 10.1: Composite GrassChopper Cost Breakdown
35
increase. The idea for using composite structures became more developed as the cost difference between an
aluminum and composite helicopter became more significant.
An important note in Figure 10.2 is that in April/May there are two data points per date. The higher
value is that of the aluminum model and the lower for the composite model. The final cost for the aluminum
model is $275K and $261K for the composite model.
100
150
200
250
300
350
400
450
5-Sep-05 19-Nov-05 2-Feb-06 18-Apr-06 2-Jul-06Date
Cost
($10
00)
In comparison to the piston-engine
equivalent trainers, the GrassChopper is
approximately $50K more expensive, as seen in
Figure 10.3. Note, however, the GrassChopper
out performs the Robinson R22 in maximum
velocity, maximum range, and maximum
endurance as discussed in section 7.3 of this
proposal. Therefore a pilot can gain experience
flying a turbine powered rotorcraft that is cost
competitive to older, out performed, piston
helicopters.
PSU250 Used
Figure 10.2: Record of the Cost for the GrassChopper
10.4 – Direct Operating Cost
The direct operating cost (DOC) of any aircraft consists of three main contributors: fuel and oil
consumption, scheduled and unscheduled maintenance, and airframe and engine overhaul. Fuel is consumed
at 0.8 lb/hp-hr which at $4.00 per gallon yields a cost of $19.20 per hour. From engine analysis, the oil
consumption contributes to approximately $0.37 per hour. Based on the analysis of other aircraft, the
airframe and engine overhaul would add approximately $50 per hour of flight do the operating cost of the
aircraft. Another important factor is the cost of the flight instructor. Assuming a helicopter pilot earns
approximately $20 per hour, this yields an approximate direct operating cost of $79.20 per hour.
Cost Comparison
$-
$50.00
$100.00$150.00
$200.00
$250.00
$300.00
Robinson R22 Schweizer300C
GrassChopper GrassChopper(composite)
Rotorcraft Model
Cos
t ($1
000)
Figure 10.3 : Comparison of Leading Piston Engine
Trainers to the GrassChopper
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38
Appendix A – MIL-STD-1374 Weight Statement MIL-STD-1374 GROUP WEIGHT STATEMENT PSU-TECHNION AIRCRAFT GRASSCHOPPER MAY 30 2006 (INCLUDING ROTORCRAFT) PAGE 1 ESTIMATED CALCULATED ACTUAL (CROSS OUT THOSE NOT APPLICABLE)
1 WING GROUP WINGLETS GLOVE / LEX WING 2 TOTAL 3 BASIC STRUCTURE 4 CENTER SECTION 5 INTERMEDIATE PANEL 6 OUTER PANEL 7 SECONDARY STRUCTURE 8 AILERONS / ELEVONS 9 SPOILERS NONE
10 FLAPS - TRAILING EDGE NONE 11 - LEADING EDGE NONE 12 SLATS 13 14 15 ROTOR GROUP 98.47 16 BLADE ASSEMBLY 17 HUB & HINGE 1.8 18 19 EMPENNAGE GROUP CANARD HORIZ. STAB. VERT FIN TAILROTOR 20 TOTAL NONE 1.8 4.5 6.3 21 BASIC STRUCTURE 22 SECONDARY STRUCTURE 23 CONTROL SURFACES 24 ( INCL. BALANCE WEIGHTS ) 25 BLADES 26 HUB & HINGE 27 ROTOR / FAN DUCT & ROTOR SUPTS 28 29 30 FUSELAGE GROUP FUS / HULL BOOMS 31 TOTAL 171.6 32 BASIC STRUCTURE 33 SECONDARY STRUCTURE 34 ENCLOSURES, FLOORING, ETC. 35 DOORS, RAMPS, PANELS & MISC. 36 37 38 ALIGHTING GEAR GROUP 39 TOTAL 126.2 40 RUNNING GEAR / FLOATS / SKIS 41 STRUCTURE 42 CONTROLS 43 44 45 ENGINE SECTION OR NACELLE GROUP 46 LOCATION 47 TOTAL – EACH LOCATION 48 49 50 51 AIR INDUCTION GROUP 52 LOCATION 53 TOTAL – EACH LOCATION 20 54 INLETS 55 DUCTS, ETC.
39
MIL-STD-1374 PSU-TECHNION GRASSCHOPPER MAY 30 2006 PAGE 2
56 TOTAL STRUCTURE 422.57 57 58 PROPULSION GROUP 132.2 59 ENGINE 105.2 60 ENGINE INSTALLATION 61 ACCESSORY GEAR BOXES & DRIVE 27 62 EXHAUST SYSTEM 63 ENGINE COOLING 64 WATER INJECTION 65 ENGINE CONTROLS 66 STARTING SYSTEM 67 PROPELLER / FAN INSTALLATION 68 LUBRICATING SYSTEM 69 FUEL SYSTEM 70 TANKS - PROTECTED 71 - UNPROTECTED 72 PLUMBING, ETC. 73 74 DRIVE SYSTEM 118.4 75 GEAR BOXES, LUB SYS & RTR BRK 76 TRANSMISSION DRIVE 77 ROTOR SHAFTS 78 GAS DRIVE 79 80 FLIGHT CONTROLS GROUP 133.2 81 COCKPIT CONTROLS 13.5 82 AUTOMATIC FLIGHT CONTROL SYSTEM 83 SYSTEM CONTROLS 13.5 84 AUXILIARY POWER GROUP 85 INSTRUMENTS GROUP 10.3 86 HYDRAULIC GROUP 87 PNEUMATIC GROUP 88 ELECTRICAL GROUP 19.1 89 AVIONICS GROUP 34.4 90 EQUIPMENT 91 INSTALLATION 92 ARMAMENT GROUP 93 FURNISHINGS & EQUIPMENT GROUP 32.4 94 ACCOMMODATION FOR PERSONNEL 95 MISCELLANEOUS EQUIPMENT 96 FURNISHINGS 97 EMERGENCY EQUIPMENT 98 AIR CONDITIONING GROUP 99 ANTI-ICING GROUP
100 PHOTOGRAPHIC GROUP 101 LOAD & HANDLING GROUP 102 AIRCRAFT HANDLING 103 LOAD HANDLING 104 105 TOTAL SYSTEMS AND EQUIP. 106 BALLAST GROUP 107 MANUFACTURING VARIATION 9.2 108 CONTINGENCY 34.5 109 110 TOTAL CONTRACTOR CONTROLLED 111 TOTAL GOVERNMENT FURNISHED EQUIP. 112 TOTAL CONTRACTOR - RESPONSIBLE 113 TOTAL GOVERNMENT - RESPONSIBLE 114 TOTAL WEIGHT EMPTY PG. 2-3 806.37
40
MIL-STD-1374 PSU-TECHNION GRASSCHOPPER MAY 30 2006 PAGE 3 115 LOAD CONDITION PRIMARY 116 117 WEIGHT EMPTY 806.37 118 CREW ( QTY _ 0 ) 119 UNUSABLE FUEL 120 OIL 121 TRAPPED 122 ENGINE 7 123 TRANSMISSION 7 124 AUX. FUEL TANKS QTY 125 INTERNAL 126 EXTERNAL 127 128 WATER INJECTION FLUID 129 BAGGAGE 130 GUN INSTALLATIONS 131 GUNS LOC FIX. OR FLEX. QTY CAL. NONE 132 133 134 SUPPORTS 135 WEAPONS PROVISIONS NONE 136 137 138 139 140 CHAFF ( QTY _________ ) 141 FLARES ( QTY _________ ) 142 143 144 SURVIVAL KITS 145 LIFE RAFTS NONE 146 OXYGEN NONE 147 148 149 150 OPERATING WEIGHT 151 PASSENGERS 400 152 153 CARGO 44 154 155 AMMUNITION QTY CAL. 156 157 158 WEAPONS NONE 159 160 161 162 163 164 ZERO FUEL WEIGHT 165 USABLE FUEL TYPE LOC GALS 276 166 INTERNAL 167 168 EXTERNAL 169 170 TOTAL USEFUL LOAD 734 171 GROSS WEIGHT 1540.37
