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Aircraft Design: A Conceptual Approach, Sixth Edition

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Aircraft Design: A Conceptual Approach, Sixth Edition

Aircraft Design:
A Conceptual Approach
Sixth Edition
Daniel P. Raymer, PhD.
Conceptual Research Corporation
Playa del Rey, California, USA
ISBN 978-1-60086-911-2
Copyright © 2018 by Daniel P. Raymer. All rights reserved.
Published by the AIAA Education Series
American Institute of Aeronautics and Astronautics, Inc.
1801 Alexander Bell Drive, Reston, Virginia 20191-4344
This book is copyright and is not legally available in electronic format on this or any other
internet site.
ALL SUCH ‘FREE DOWNLOAD’ COPIES ARE ILLEGAL!
Please do not download pirate copies – it is immoral and it contributes to a world in which
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The book is available for purchase at most online bookstores including Amazon, and can
also be purchased directly from the publisher, at these links.
http://www.amazon.com/exec/obidos/ASIN/1624104908/danraymesaircrde
https://arc.aiaa.org/doi/book/10.2514/4.104909
Below is the full Table of Contents followed by the Preface, which introduces the book.
For more information about this book, aircraft design short courses, the RDSwin design
software, and free downloads of other publications by the author, see the author’s website:
www.aircraftdesign.com
CONTENTS
Preface
Author's Note Concerning Use of Metric Weight Units
Nomenclature
Chapter 1 Design—A Separate Discipline
1.1 What Is Design?
1.2
Design: How Does It Start?
1.3
An Airplane Designer: How Can I Become One?
1.4 The Book: What Is Here and How It Is Organized
Chapter 2 Overview of the Design Process
2.1 Requirements
2.2
Phases of Aircraft Design
2.3
Aircraft Conceptual Design Process
2.4 Integrated Product Development and Aircraft Design
Chapter 3 Sizing from a Conceptual Sketch
3.1 Introduction
3.2 Takeoff-Weight Buildup
3.3
-
Weight Estimation
3.4
Fuel
-
Fraction Estimation
3.5
Takeoff
-
Weight Calculation
3.6 Design Example: ASW Aircraft
Chapter 4 Airfoil and Wing/Tail Geometry Selection
4.1 Introduction
4.2
Airfoil Selection
4.3 Wing Geometry
4.4
Biplane Wings
4.5
Tail Geometry and Arrangement
Chapter 5 Thrust-to-Weight Ratio and Wing Loading
5.1 Introduction
5.2
Thrust
-
to
-
Weight Ratio
5.3
Wing Loading
5.4 Selection of Thrust to Weight and Wing Loading
Chapter 6 Initial Sizing
6.1 Introduction
6.2
"Rubber" vs "Fixed
-
Size" Engines
6.3 Rubber-Engine Sizing
6.4
Fixed
-
Engine Sizing
6.5
Geometry Sizing
6.6 Control-Surface Sizing
Chapter 7 Configuration Layout and Loft
7.1 Introduction
7.2 End Products of Configuration Layout
7.3
Conic Lofting
7.4
Conic Fuselage Development
7.5
Flat
-
Wrap Fuselage Lofting
7.6
Circle
-
to
-
Square Adapter
7.7
Loft Verification via Buttock
-
Plane Cuts
7.8 Wing/Tail Layout and Loft
7.9
Wetted
-
Area Determination
7.10
Volume Determination
7.11 Use of Computer-Aided Design (CAD)
in Conceptual Design
Chapter 8 Special Considerations in Configuration Layout
8.1 Introduction
8.2 Aerodynamic Considerations
8.3
Structur
al Considerations
8.4 Radar Detectability
8.5
Infrared Detectability
8.6
Visual Detectability
8.7
Aural Signature
8.8
Vulnerability Considerations
8.9
Crashworthiness Considerations
8.10 Producibility Considerations
8.11
Maintainability Considerations
Chapter 9 Crew Station, Passengers, and Payload
9.1 Introduction
9.2
Crew Station
9.3
Passenger Compartment
9.4
Cargo Provisions
9.5
Weapons Carriage
9.6 Gun Installation
Chapter 10 Propulsion and Fuel System Integration
10.1 Introduction
10.2
Propulsion Overview and Selection
10.3
Jet
-
Engine Integration
10.4
Propeller
-
Engine Integration
10.5
Fuel System
10.6
Green Propulsion
Chapter 11 Landing Gear and Subsystems
11.1 Introduction
11.2 Landing-Gear Arrangements
11.3
Tire Sizing
11.4
Shock Absorbers
11.5 Castoring-Wheel Geometry
11.6
Gear
-
Retraction Geometry
11.7
Seaplanes
11.8 Subsystems
Intermission: Step-by-Step Development of a New Design
Chapter 12 Aerodynamics
12.1 Introduction
12.2
Aerodynamic Forces
12.3
Aerodynamic Coefficients
12.4
Lift
12.5
Parasite (Zero
-
Lift) Drag
12.6
Drag Due to Lift (Including Induced
Drag)
12.7 Computational Fluid Dynamics
Chapter 13 Propulsion
13.1 Aircraft Thrust—The Big Picture
13.2 Jet-Engine Thrust Considerations
13.3
Jet
-
Engine Installed Thrust
13.4
Part Power Operation
13.5
Piston
-
Engine Overvi
ew
13.6 Propeller Analysis
13.7
Piston
-
Prop Thrust
Corrections
13.8 Turboprop Performance
Chapter 14 Structures and Loads
14.1 Introduction
14.2 Loads Categories
14.3
Air Loads
14.4
Inertial Loads
14.5
Powerplan
t Loads
14.6
Landing
-
Gear Loads
14.7
Structures Fundamentals
14.8
Material Selection
14.9
Material Properties
14.10
Structural
-
Analysis Fundamentals
14.11 Finite Element Structural Analysis
Chapter 15 Weights
15.1 Introduction
15.2 Approximate Weight Methods
15.3
Aircraft Statistical Weights Method
15.4 Additional Considerations in Weights Estimation
Chapter 16 Stability, Control, and Handling Qualities
16.1 Introduction
16.2
Coordinate S
ystems and Definitions
16.3
Longitudinal Static Stability and Control
16.4
Lateral
-
Directional Static Stability and Control
16.5
Stick
-
Free Stability
16.6
Effects of Flexibility
16.7
Dynamic Stability
16.8
Quasi Steady State
16.9
Inertia Coupling
16.10 Handling Qualities
Chapter 17 Performance and Flight Mechanics
17.1 Introduction and Equations of Motion
17.2 Steady Level Flight
17.3
Steady Climbing and Descending Flight
17.4
Level Turning Fl
ight
17.5
Gliding
Flight
17.6
Energy
-
Maneuverability Methods
17.7
Operating Envelope
17.8
Takeoff Analysis
17.9
Landing Analysis
17.10 Other Fighter Performance Measures of Merit
Chapter 18 Cost Analysis
18.1 Introduction
18.2
Elements of Life
-
Cycle Cost
18.3
Cost
-
Estimating Methods
18.4
RDT&E and Production Costs
18.5
Operations and Maintenance Costs
18.6
Cost Measures of Merit (Military)
18.7 Aircraft and Airline Economics
Chapter 19 Sizing and Trade Studies
19.1 Introduction
19.2
Detailed Sizing Methods
19.3
Improved Conceptual Sizing Methods
19.4
Classic Optimization
Sizing Matrix and Carpet Plots
19.5 Trade Studies
Chapter 20 Electric Aircraft
20.1 Introduction
20.2 Review of Physics & Units
20.3 Why Spark?
20.4 Electric Motor Basics
20.5 Power Supply: Batteries
20.6 Power Supply: Fuel Cells
20.7 Power Supply: Hybrid-Electric
20.8 Power Supply: Solar Cells
20.9 Power Supply: Beamed Power
20.10 Electric Aircraft Run-Time, Range, Loiter, and Climb
20.11 Electric Aircraft Initial Sizing
Chapter
2
1
Vertical Flight
Jet and Prop
21.1 Introduction
2
1
.2
Jet VTOL
21.3 Prop VTOL and Helicopter
Chapter 22 Extremes of Flight
22.1 Introduction
2
2
.2
Rockets, Launch Vehicles, and Spacecraft
2
2
.3
Hypersonic Vehicles
22.4 Lighter Than Air
Chapter 23 Design of Unique Aircraft Concepts
23.1 Introduction
23.2 Flying Wing, Lifting Fuselage, and Blended Wing Body
23.
3
Delta and Double
-
Delta Wing
23.
4
Forward
-
Swept Wing
23.
5
Canard
-
Pusher
23.6 Multifuselage
23.
7
Asymmetric Airplanes
23.
8
Joined Wing
23.
9
Some More Innovative Wings
23.
10
Wing
-
in
-
Ground
-
Effect
23.11 Unmanned/Uninhabited Aircraft
23.12 Derivative Aircraft Design
Chapter 24 Conceptual Design Examples
24.1 Introduction
2
4
.2
Single
-
Seat Aerobatic Homebuilt
24.3 Lightweight Supercruise Fighter
Appendix A Unit Conversion
Appendix B Standard Atmosphere
Appendix C Airspeed
Appendix D Airfoil Data
Appendix E Typical Engine Performance Curves
E.1 Afterburning Turbofan
E.2
High
-
Bypass Turbofan
E.3 Turboprop
Appendix F Design Requirements and Specifications
Questions
References
Index
Supporting Materials
PREFACE
Aircraft Design is a challenging, rewarding, and fun career. There are dozens of
different activities involved in creating a new air vehicle concept, different specialties
ranging from initial configuration layout to system optimization and cost estimation.
These activities can be grouped into two equally important aspects of aircraft design:
design layout and design analysis. While some people do both, in most cases these
differing aspects attract different types of people. Certain people love playing with
numbers and computers, whereas others can't stop doodling on every piece of paper
within reach.
This book offers a balanced overview of these two aspects of design, integrated
together and presented in the manner typically seen in an aircraft design project at a
major aerospace company. Whichever aspect you may lean towards, the book should help
get you started and will provide a resource of material throughout your career.
Aircraft design depends on the reliable calculation of numbers but in the end, the
only thing that actually gets built is the configuration concept shown on the drawing or
CAD file. Its creation is not a trivial task of drafting based upon the analysis results, but
rather it is a key element of the overall design process and ultimately determines the
performance, weight, and cost of the aircraft. Bluntly stated, if you don't have a good
drawing, you don't have an aircraft design. The “Conceptual Approach” mentioned in the
book’s title refers to a design process centered around a realistic concept layout.
It is difficult to visualize and draw a new aircraft that has a streamlined
aerodynamic shape and an efficient internal layout and yet satisfies an incredible number
of real-world constraints and design specifications. Aircraft conceptual design layout is a
rare talent that takes years to cultivate. Although to some extent good designers are
“born, not made,” the proven methods and best practices of aircraft configuration layout
can be taught and are covered here in the first half of this book. These apply equally to
traditional drafting table drawings and to modern computer-aided design.
It is also true that a nice aircraft drawing is nothing without the analytical results
to support it, and it will be a much better design if clever optimization methods are
employed. So, a good designer or design team must find an appropriate balance between
design layout and design analysis. The second half of this book covers analysis and
optimization methods that will tell you if the design works, if it meets its design
requirements, and how you can make it better in the next drawing.
Writing – and rewriting - this book has been an educating and humbling
experience. It is my sincere wish that it helps aspiring aircraft designers to “learn the
ropes” more quickly. My greatest pride in the previous editions has been the thanks from
the students who've used the book in their design classes, and the designers of built-and-
flown airplanes who've told me that they made extensive use of my book. Thanks—that
means a lot.
Daniel P. Raymer
Los Angeles, California
June 2018
... Building the landing gear and its wheel disc brakes is another challenging aspect of this work. This paper referred to the standard equations found in Daniel Raymer's book on airplane design [5]. Extracted linkages include those between wheel diameter and thickness, wheel height and thickness, and wheel design versus disc brake design parameters. ...
... Extracted linkages include those between wheel diameter and thickness, wheel height and thickness, and wheel design versus disc brake design parameters. Daniel Raymer [5] compiled the formulas used to build the disc brake used in this study. ...
... Under the design process, the determination of height, mass, volume, and hollow section-based inner and outer diameters of the landing stick are executed. One of the major factors involved in the design process is the fineness ratio (FR), which mostly lies between 5 and 10 [5]. ...
Article
Full-text available
Citation: Raja, V.; Gnanasekaran, R.K.; Rajendran, P.; Mohd Ali, A.; Rasheed, R.; AL-bonsrulah, H.A.Z.; Al-Bahrani, M. Asymmetrical Damage Aspects Based Investigations on the Disc Brake of Long-Range UAVs through Verified Computational Coupled Approaches.
... It persists because of the influence of the upstream boundary layer. Third, the drag caused by the formation of shocks at supersonic and high supersonic speeds is referred to as wave drag [1]. First, the flow past a projectile is an example of external flow. ...
... Specifically, with a cavity, the literation shows only a few studies, based on the experimental work. Figure 2: Sketch of a multi-step afterbody [11] Therefore, this research aims to reproduce/validate the Raymer [1] work using the advanced soft computing tool ANSYS fluent-based CFD. ...
... Fig. 3 depicts an axisymmetric convergent-divergent nozzle model connected to a concentrated axisymmetric duct with annular rectangular cavities. The dimensions for the convergent-divergent nozzle with suddenly expanded circular are based on an experimental setup by Raymer [1]. ...
Article
Full-text available
In this study, a duct is considered and special attention is paid to a passive method for the control of the base pressure relying on the use of a cavity with a variable aspect ratio. The Mach number considered is 1.8, and the area ratio of the duct is 2.56. In particular, two cavities are examined, their sizes being 3:3 and 6:3. The used L/D spans the interval 1-10 while the NPRs (nozzle pressure ratio) range from 2 to 9. The results show that the control becomes effective once the nozzles are correctly expanded or under-expanded. The pressure contours at different NPR and L/D are presented. It is shown that the NPR and cavity location strongly influence the base pressure. The NPR, Mach number, and cavity aspect ratio have a strong effect on the base pressure in the wake region. KEYWORDS Base pressure; passive control; computational fluid dynamics (CFD); supersonic; aerodynamics; cavity 1 Introduction Baseflow studies at high Mach numbers continue to be an area of research because of their many different applications in the aerospace industry, defence, and space research. The flow field at a blunt base is characterized by flow separation, and reattachment can be divided into two regions: the central and the separated regions. The shear layer will generate strong vortices in the wake region, increasing the overall drag force. The net drag force comprises skin friction, wave, and base drag for any aerodynamic body. For the Mach number, less than unity total drag will consist of skin friction drag and base drag. However, wave drag will be an additional component at sonic and supersonic Mach numbers. The literature review reveals that the base drag component is two-thirds of the total drag in transonic speed.
... The aircraft aerodynamic database is built by integrating simple and reliable semiempirical formulation like those proposed by Roskam [29], Raymer [30] and Sforza [31]. ...
Article
Full-text available
Structural health monitoring represents an interesting enabling technology towards increasing aviation safety and reducing operating costs by unlocking novel maintenance approaches and procedures. However, the benefits of such a technology are limited to maintenance costs reductions by cutting or even eliminating some maintenance scheduled checks. The key limitation to move a step further in exploiting structural health monitoring technology is represented by the regulation imposed in sizing aircraft composite structures. A safety margin of 2.0 is usually applied to estimate the ultimate loading that composite structures must withstand. This limitation is imposed since physical nondestructive inspection of composite structures is really challenging or even impossible in some cases. However, a structural health monitoring system represents a viable way for a real time check for the health status of a composite structure. Thus, the introduction of structural health monitoring should help into reducing the stringent safety margin imposed by aviation regulation for a safe design of composite structures. By assuming a safety margin reduction from 2.0 to 1.75 thanks to the installation of permanently attached sensors for structural health diagnostics, this paper assesses the potential fuel savings and direct operating costs through a multidisciplinary analysis on a A220-like aircraft. According to the foreseen level of technology, addressed through the number of sensors per square meter, a DOC saving from 2% up to 5% is achievable preserving, at the same time, all the key aircraft performance.
... The fuselage is designed using an inside-out approach, following the passenger capacity requirements. The wing planform results from empirical rules [25,26]. For the empennage sizing, we assume statistical tail volume coefficients from reference aircraft data. ...
Conference Paper
Full-text available
The aircraft’s environmental performance on fleet level is so far completely decoupled from the design process. The climate impact from aviation arising from non-CO2 effects are largely independent from CO2 emissions, but rather depend on the atmospheric state. Previously complex climate-chemistry models were used to evaluate the non-CO2 emissions impact on climate. This is far too computationally demanding for a multidisciplinary design optimisation (MDO) process, requiring a multitude of climate impact evaluations. The question then is, how to efficiently design the next generation climate optimal aircraft? In this paper, a new concept for designing aircraft with minimum climate impact using Climate Functions for Aircraft Design (CFAD) is presented. The content of this paper provides an overview of the development of these innovative CFAD and demonstrates the ability to be integrated in an existing MDO framework. The miti- gation potential by optimising aircraft design using CFAD is analysed with respect to different cruise conditions and by minimizing the overall climate impact. To validate the CFAD, a higher fidelity assessment is carried out. Finally, the key performance indicators, i.e. fuel consumption, flight time and operating cost, of the optimised aircraft design are compared to that of the reference aircraft.
... The mass of the BLI propulsor was estimated using the built-in method in Initiator for turboprop engines, which is uses a database of known turboprop engines to predict bare engine mass. The APPU engine cycle efficiency was predicted using a statistical model from Raymer [16], based on known turboshaft engines. The benefit of BLI to propulsive efficiency was calculated based on an actuator disk analogy in potential flow. ...
Conference Paper
Full-text available
The concept of an "Auxiliary Power and Propulsion Unit" (APPU) is introduced, which consists of a Boundary-layer-ingesting (BLI) propulsor with an engine mounted at the rear of an passenger aircraft fuselage, replacing the Auxiliary Power Unit (APU) and contributing around 10% of total cruise thrust, as well as auxiliary power. This APPU unit is using hydrogen provided by an additional tank installed in the tailcone of the aircraft. The concept is aimed at lowering the threshold to installing both hydrogen-driven propulsion and BLI propulsors on aircraft in the short term, while minimizing resulting operational risk. The concept has been investigated using a preliminary aircraft synthesis tool and further component-level mass estimates. Operational aspects, sensitivities and limits to the design have been investigated. Estimates of mission fuel burn find that CO 2 emissions emissions reduce roughly proportionally to the APPU thrust share, with additional savings due to improved overall efficiency. Further improvements are deemed feasible and are the topic of ongoing research.
... The landing gear mass is estimated using Raymer's method [43], because this method makes the sizing dependent on the aircraft landing mass and the landing gear height, two parameters which will sensibly differ between the kerosene and the LH 2 aircraft versions. ...
Article
Full-text available
Zero-carbon-dioxide-emitting hydrogen-powered aircraft have, in recent decades, come back on the stage as promising protagonists in the fight against global warming. The main cause for the reduced performance of liquid hydrogen aircraft lays in the fuel storage, which demands the use of voluminous and heavy tanks. Literature on the topic shows that the optimal fuel storage solution depends on the aircraft range category, but most studies disagree on which solution is optimal for each category. The objective of this research was to identify and compare possible solutions to the integration of the hydrogen fuel containment system on regional, short/medium- and large passenger aircraft, and to understand why and how the optimal tank integration strategy depends on the aircraft category. This objective was pursued by creating a design and analysis framework for CS-25 aircraft capable of appreciating the effects that different combinations of tank structure, fuselage diameter, tank layout, shape, venting pressure and pressure control generate at aircraft level. Despite that no large differences among categories were found, the following main observations were made: (1) using an integral tank structure was found to be increasingly more beneficial with increasing aircraft range/size. (2) The use of a forward tank in combination with the aft one appeared to be always beneficial in terms of energy consumption. (3) The increase in fuselage diameter is detrimental, especially when an extra aisle is not required and a double-deck cabin is not feasible. (4) Direct venting has, when done efficiently, a small positive effect. (5) The optimal venting pressure varies with the aircraft configuration, performance, and mission. The impact on performance from sizing the tank for missions longer than the harmonic one was also quantified.
... The trim solver is obtained, for a given set of kinematic variables, x, and commands, u, by combining the aerodynamic/propulsive loads (X, Y, Z, L, M, N) produced by each vehicle component with the gravitational and inertial loads. The aerodynamic interaction effects between components are either neglected or modeled through semi-empirical coefficients [16]. For each rotor, the aeroelastic response to four commands (namely, angular speed Ω, collective pitch θ 0 , lateral pitch θ c and longitudinal pitch θ c ) and to the hub kinematics (given in terms of u, v , w , p, q, r ) is evaluated. ...
Conference Paper
The aim of the proposed paper is the development of a general approach capable of determining the set of trim commands, aerodynamic controls surfaces actuation and/or engine thrust regulation, that allow specific steady-state flight conditions of non-conventional VTOLs, under given performance constraints. Specifically, an energy-based trim algorithm is introduced, based on the minimisation of a cost function. This approach is particularly interesting in that, in case of multi-rotor system, it allows to determine the control settings such to guarantee the required flight condition while using the redundant controls to optimise selected target functions, as for instance performance or noise emission. The proposed approach will be applied to a quadcopter configuration, for validation purposes, and a hexacopter configuration for which a comparison with a conventional trim procedure will be performed. Three different control strategies are applied: a pure rotor angular velocity control, a pure blade collective pitch control and a combined angular velocity and blade collective pitch control.
Chapter
Estimation of aircraft weight in flight, which is a dominant parameter related to its flight performance, was studied. A typical approach for the estimation is to use a simple combination of initial weight on the ground and fuel consumption in air obtained with fuel tank gauge or by accumulating fuel flow. This approach is insufficient when the flight performance will be estimated as accurately as possible because some of the available measurements are not utilized. Therefore, the forward–backward smoother derived from Kalman filter was applied to the estimation with our revision of the smoother to fuse terminal weight on the ground additionally. According to numerical simulations, our method estimated the weight in smaller errors than not only the typical approach but also a fixed-interval smoother, which is used generally for an off-line estimation problem. Moreover, application to actual flight data showed that our method improved the estimated standard deviation by approximately three percent at maximum compared to the fixed-interval smoother.KeywordsWeight estimationSensor fusionForward–backward smoother
Chapter
The addition of several power sources in an aircraft increases the complexity of the sizing and energy management problem while allowing a system redundancy that makes aircrafts safer. The optimization of the sizing and the energy management of hybrid electric aircraft powertrains can be accomplished using comprehensive mathematical models from the aircraft and its power sources, reducing the load of experimental activities that turn to be expensive and time-consuming. In this work, the authors apply an optimization method to obtain two optimized energy management strategies to be applied to two different types of all-electric aircraft: a general aviation powertrain and an electric vertical take-off and landing powertrain. These two aircrafts are designed to employ the same power sources configuration with a hydrogen-fueled fuel cell and a battery pack. The energy management optimization was performed to maximize the traveled distance while keeping the battery’s state of charge difference at a minimum, observing the power sources restrictions. In addition, for the second powertrain, the optimization of the power sources was performed. The analysis of the results shows that using the proposed method, the general aviation powertrain improves the traveled distance by 2.78%, reducing the equivalent energy consumption by 2.73%, and the electric vertical take-off and landing powertrain reduces the equivalent power consumption and guarantees the same battery’s state of charge at the start and at the end of the flight allowing a non-plugin operation.
Article
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The paper presents the process of designing an unmanned Micro class aircraft, from the analysis of the dynamically developing market and the condition of the Polish Armed Forces to construction of objects and flight test. The possibilities and limitations of using miniature UAVs on the modern battlefield were determined. For the designed UAV the propulsion was selected based on tests carried out on the engine test bench. The avionics equipment was selected based on components readily available on the market. The object was then made and inspected in flight. During the flight tests, the aircraft performance was verified and compared with the assumptions. It has been shown that the developed object is able to fulfill the reconnaissance tasks entrusted to it, while maintaining the assumed simplicity of construction and low cost of execution and service.
ResearchGate has not been able to resolve any references for this publication.