This document discusses methods for estimating aerodynamic properties of aircraft parts and surfaces needed to determine stability and control characteristics. These include estimating lift curve slope, zero-lift angle of attack, pitch moment, and downwash parameter for wings, horizontal tails, and vertical tails through relationships involving airfoil properties, geometry, and flow characteristics. Semi-empirical equations are presented to calculate these properties for straight tapered wings and tails using parameters like aspect ratio, taper ratio, sweep, and position.
Airfoil properties, shapes & structural dynamical features are described. Nomenclature or the classification types are presented along with the application.
Common methods for analysis of the structural dynamics on a wing or blade are presented along with the possible applications.
ME 438 Aerodynamics is a course taught by Dr. Bilal Siddiqui at DHA Suffa University. This set of lectures start from the basic and all the way to aerodynamic coefficients and center of pressure variations with angle of attack.
The flow across an airfoil is studied for different angle of attack. The CFD analysis results are documented and studied for different angle of attack using fluent & gambit.
Airfoil properties, shapes & structural dynamical features are described. Nomenclature or the classification types are presented along with the application.
Common methods for analysis of the structural dynamics on a wing or blade are presented along with the possible applications.
ME 438 Aerodynamics is a course taught by Dr. Bilal Siddiqui at DHA Suffa University. This set of lectures start from the basic and all the way to aerodynamic coefficients and center of pressure variations with angle of attack.
The flow across an airfoil is studied for different angle of attack. The CFD analysis results are documented and studied for different angle of attack using fluent & gambit.
Design Analysis Of Uav (Unmanned Air Vehicle) Using NACA 0012 Aerofoil ProfileDr. Bhuiyan S. M. Ebna Hai
This research work is concerned with the application of conceptual design of Unmanned Air Vehicle (UAV). UAV is used for surveillance and reconnaissance to serve for the defense as well as national security and intelligence purpose. Here NACA 0012 aerofoil profile is used to design UAV by using CFD (Computational Fluid Dynamics) software. The aim of this research is to investigate the flow patterns and determine the aerodynamic characteristics of NACA 0012 profile by varying the angle of attack and Reynolds Number numerically. The research is carried out with symmetric aerofoil with the chord length of 0.1m. The research work explained different aerodynamic characteristics like lift force and drag force, lift and drag coefficient, pressure distribution over aerofoil etc.
Dynamic aerodynamic structural coupling numerical simulation on the flexible ...ijmech
Most biological flyers undergo orderly deformation in flight, and the deformations of wings lead to complex fluid-structure interactions. In this paper, an aerodynamic-structural coupling method of flapping wing is developed based on ANSYS to simulate the flapping of flexible wing.
Firstly, a three-dimensional model of the cicada’s wing is established. Then, numerical simulation method of unsteady flow field and structure calculation method of nonlinear large deformation are studied basing on Fluent module and Transient Structural module in ANSYS,
the examples are used to prove the validity of the method. Finally, Fluent module and Transient Structural module are connected through the System Coupling module to make a two-way fluidstructure Coupling computational framework. Comparing with the rigid wing of a cicada, the
coupling results of the flexible wing shows that the flexible deformation can increase the aerodynamic performances of flapping flight.
Strategic design of aircraft wings have evolved over time for maximum fuel efficiency. One of such ideas involves winglet which has been known
to reduce turbulence at the tip of the wings. This study intends to investigate the
differences in drag and lift forces generated at aeroplane wings with and without winglet at cruising speed using FEM. Simulations were performed in the
SST turbulence model of CFD and the results are compared to that of the experimental and theoretical models. The simulation showed that the lift increased
by 26.0% and the drag decreased by 74.6% for the winglet at cruising speed.
Notes on the formula for minimum horizontal radius. Rev. 05 (January 2016) - Added notes on different measures of speed and different speeds for horizontal and vertical
design. Added row h to values table. Modified layout. Added extra note and reference on 3D road design.
Angle of Attack | Q & A | Question Analysis | Flight Mechanics | GATE AerospaceAge of Aerospace
Question Analysis, Book Reference, Important Concepts, and topic wise Solutions for the topic "Angle of Attack" are time-stamped below. Access the study materials, presentation, links to previous and next lectures and further information in the description section.
30 note - sems and measure of the total elastic envelope (ee)Miguel Cabral Martín
.0 - NOTE - SEMS and MEASURE of the TOTAL ELASTIC ENVELOPE (EE)
In the next Fig. 1 we suppose the white curve AB is the structural axis of a wing which represents its elastic envelope (EE). The lines AC and BC represent the tangents of the rotation angles of points A and B.
Improving the Hydraulic Efficiency of Centrifugal Pumps through Computational...IJERA Editor
The design and optimization of turbo machine impellers such as those in pumps and turbines is a highly complicated task due to the complex three-dimensional shape of the impeller blades and surrounding devices. Small differences in geometry can lead to significant changes in the performance of these machines. We report here an efficient numerical technique that automatically optimizes the geometry of these blades for maximum performance. The technique combines, mathematical modeling of the impeller blades using non-uniform rational B-spline (NURBS), Computational fluid dynamics (CFD) with Geometry Parameterizations in turbulent flow simulation and the Globalized and bounded Nelder-Mead (GBNM) algorithm in geometry optimization.
For Video Lecture of this presentation: https://youtu.be/NAjezfbWh4Y
The topics covered in this session are, drag, categories of drag, drag polar equation and drag polar graph, drag polar derivation, induced drag coefficient.
Attention! "Gate Aerospace Engineering aspirants", A virtual guide for gate aerospace engineering is provided in "Age of Aerospace" blog for helping you meticulously prepare for gate examination. Respective notes of individual subjects are provided as 'Embedded Google Docs' which are frequently updated. This comprehensive guide is intended to efficiently serve as an extensive collection of online resources for "GATE Aerospace Engineering" which can be accessed free of cost. Use the following link to access the study material
https://ageofaerospace.blogspot.com/p/gate-aerospace.html
Design Analysis Of Uav (Unmanned Air Vehicle) Using NACA 0012 Aerofoil ProfileDr. Bhuiyan S. M. Ebna Hai
This research work is concerned with the application of conceptual design of Unmanned Air Vehicle (UAV). UAV is used for surveillance and reconnaissance to serve for the defense as well as national security and intelligence purpose. Here NACA 0012 aerofoil profile is used to design UAV by using CFD (Computational Fluid Dynamics) software. The aim of this research is to investigate the flow patterns and determine the aerodynamic characteristics of NACA 0012 profile by varying the angle of attack and Reynolds Number numerically. The research is carried out with symmetric aerofoil with the chord length of 0.1m. The research work explained different aerodynamic characteristics like lift force and drag force, lift and drag coefficient, pressure distribution over aerofoil etc.
Dynamic aerodynamic structural coupling numerical simulation on the flexible ...ijmech
Most biological flyers undergo orderly deformation in flight, and the deformations of wings lead to complex fluid-structure interactions. In this paper, an aerodynamic-structural coupling method of flapping wing is developed based on ANSYS to simulate the flapping of flexible wing.
Firstly, a three-dimensional model of the cicada’s wing is established. Then, numerical simulation method of unsteady flow field and structure calculation method of nonlinear large deformation are studied basing on Fluent module and Transient Structural module in ANSYS,
the examples are used to prove the validity of the method. Finally, Fluent module and Transient Structural module are connected through the System Coupling module to make a two-way fluidstructure Coupling computational framework. Comparing with the rigid wing of a cicada, the
coupling results of the flexible wing shows that the flexible deformation can increase the aerodynamic performances of flapping flight.
Strategic design of aircraft wings have evolved over time for maximum fuel efficiency. One of such ideas involves winglet which has been known
to reduce turbulence at the tip of the wings. This study intends to investigate the
differences in drag and lift forces generated at aeroplane wings with and without winglet at cruising speed using FEM. Simulations were performed in the
SST turbulence model of CFD and the results are compared to that of the experimental and theoretical models. The simulation showed that the lift increased
by 26.0% and the drag decreased by 74.6% for the winglet at cruising speed.
Notes on the formula for minimum horizontal radius. Rev. 05 (January 2016) - Added notes on different measures of speed and different speeds for horizontal and vertical
design. Added row h to values table. Modified layout. Added extra note and reference on 3D road design.
Angle of Attack | Q & A | Question Analysis | Flight Mechanics | GATE AerospaceAge of Aerospace
Question Analysis, Book Reference, Important Concepts, and topic wise Solutions for the topic "Angle of Attack" are time-stamped below. Access the study materials, presentation, links to previous and next lectures and further information in the description section.
30 note - sems and measure of the total elastic envelope (ee)Miguel Cabral Martín
.0 - NOTE - SEMS and MEASURE of the TOTAL ELASTIC ENVELOPE (EE)
In the next Fig. 1 we suppose the white curve AB is the structural axis of a wing which represents its elastic envelope (EE). The lines AC and BC represent the tangents of the rotation angles of points A and B.
Improving the Hydraulic Efficiency of Centrifugal Pumps through Computational...IJERA Editor
The design and optimization of turbo machine impellers such as those in pumps and turbines is a highly complicated task due to the complex three-dimensional shape of the impeller blades and surrounding devices. Small differences in geometry can lead to significant changes in the performance of these machines. We report here an efficient numerical technique that automatically optimizes the geometry of these blades for maximum performance. The technique combines, mathematical modeling of the impeller blades using non-uniform rational B-spline (NURBS), Computational fluid dynamics (CFD) with Geometry Parameterizations in turbulent flow simulation and the Globalized and bounded Nelder-Mead (GBNM) algorithm in geometry optimization.
For Video Lecture of this presentation: https://youtu.be/NAjezfbWh4Y
The topics covered in this session are, drag, categories of drag, drag polar equation and drag polar graph, drag polar derivation, induced drag coefficient.
Attention! "Gate Aerospace Engineering aspirants", A virtual guide for gate aerospace engineering is provided in "Age of Aerospace" blog for helping you meticulously prepare for gate examination. Respective notes of individual subjects are provided as 'Embedded Google Docs' which are frequently updated. This comprehensive guide is intended to efficiently serve as an extensive collection of online resources for "GATE Aerospace Engineering" which can be accessed free of cost. Use the following link to access the study material
https://ageofaerospace.blogspot.com/p/gate-aerospace.html
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
The flow velocity over airfoils increases over the convex surface resulting in lower average pressure on the 'suction' side of the airfoil compared with the concave or 'pressure' side of the airfoil.
Meanwhile, viscous friction between the air and the airfoil surface slows the airflow to some extent next to the surface.
Airflow over an airfoil produces a distribution of forces over the airfoil surface.
М.Г.Гоман, А.В.Храмцовский, Е.Н.Колесников «Оценка маневренных характеристик самолета посредством численного определения областей асимптотической устойчивости», Journal of Guidance, Control, and Dynamics, Vol. 31, No. 2 (2008), стр. 329-339.
Маневренные характеристики самолета на установившихся режимах полёта обычно оцениваются на основе установившихся решений уравнений движения самолёта как твёрдого тела. В статье описан систематический метод расчета установившихся режимов полета для класса спиральных траекторий.
Mikhail G. Goman, Andrew V. Khramtsovsky, and Evgeny N. Kolesnikov "Evaluation of Aircraft Performance and Maneuverability by Computation of Attainable Equilibrium Sets", Journal of Guidance, Control, and Dynamics, Vol. 31, No. 2 (2008), pp. 329-339.
An aircraft's performance and maneuvering capabilities in steady flight conditions are usually analyzed considering the steady-states of the rigid body equations of motion. A systematic way of computation of the set of all attainable steady states for a general class of helical trajectories is presented. The proposed reconstruction of attainable equilibrium states and their local stability maps provides a comprehensive and consistent representation of the aircraft flight and maneuvering envelopes. The numerical procedure is outlined and computational examples of attainable equilibrium sets in the form of two-dimensional cross-sections of steady-state maneuver parameters are presented for three different aircraft models.
HCR's Infinite or Convergence Series (Calculations of Volume, Surface Area of...Harish Chandra Rajpoot
Three series had been derived by the author, using double-integration in polar co-ordinates, binomial expansion and β & γ-functions for determining the volume, surface-area & perimeter of elliptical-section of oblique frustum of a right circular cone. All these three series are in form of discrete summation of infinite terms which converge into finite values hence these were also named as HCR’s convergence series. These are extremely useful in case studies & practical computations.
International Refereed Journal of Engineering and Science (IRJES)irjes
International Refereed Journal of Engineering and Science (IRJES) is a leading international journal for publication of new ideas, the state of the art research results and fundamental advances in all aspects of Engineering and Science. IRJES is a open access, peer reviewed international journal with a primary objective to provide the academic community and industry for the submission of half of original research and applications
International Refereed Journal of Engineering and Science (IRJES)irjes
International Refereed Journal of Engineering and Science (IRJES) is a leading international journal for publication of new ideas, the state of the art research results and fundamental advances in all aspects of Engineering and Science. IRJES is a open access, peer reviewed international journal with a primary objective to provide the academic community and industry for the submission of half of original research and applications
Modeling and Structural Analysis of a Wing [FSI ANSYS&MATLAB] BahaaIbrahim10
In our study, analyzing aircraft’s wing with the old assumptions will not give an exact solution but
this solution (total deformation) changes according to the geometry of the cross-section of the beam, so
the total deformation of the beam may be greater or lower than the exact solution. In these two cases,
the solution is not acceptable as in the first case which the deformation is greater than the exact solution
will make more weight and cost, and Engineers design aircraft at minimum weight and less cost. But in
the second case which will make lower deformation than exact solution will be much risky as the aircraft
could fail at any time, and this case much dangerous because it threatens the life of people.
Similar to 3 estimating aerodynamic properties 9 (20)
You could be a professional graphic designer and still make mistakes. There is always the possibility of human error. On the other hand if you’re not a designer, the chances of making some common graphic design mistakes are even higher. Because you don’t know what you don’t know. That’s where this blog comes in. To make your job easier and help you create better designs, we have put together a list of common graphic design mistakes that you need to avoid.
Can AI do good? at 'offtheCanvas' India HCI preludeAlan Dix
Invited talk at 'offtheCanvas' IndiaHCI prelude, 29th June 2024.
https://www.alandix.com/academic/talks/offtheCanvas-IndiaHCI2024/
The world is being changed fundamentally by AI and we are constantly faced with newspaper headlines about its harmful effects. However, there is also the potential to both ameliorate theses harms and use the new abilities of AI to transform society for the good. Can you make the difference?
1. Stability and Control
Estimating Aerodynamic Properties
A necessary ingredient for determining the aerodynamic properties of an aircraft is to be
able to determine the aerodynamic properties of parts of the aircraft. If we look at the expression
for the pitch-moment curve slope, we can see some of the parameters that need to be estimated:
(1)
Here we see that in order to evaluate this expression we need to estimate .
Furthermore, these same parameters show up in the other coefficients as well. If we look at the
control effectiveness and control power terms, the additional parameter, , must be estimated.
For the zero-lift pitch-moment and the zero lift angle-of-attack, additional parameters such as the
wing-body zero-lift pitch-moment, and the wing-body zero-lift angle-of-attack are needed. Here
we will examine some ways of estimating some of these quantities. The primary focus will be on
the first four parameters identified above.
The discussion will be limited to straight tapered wings (and tails). A straight tapered
surface is one where the leading and trailing edges are straight. Most wings fall into this
category.
Wing Section Properties
If we slice a wing parallel to the free stream velocity the cross section that is visible is
the wing stream-wise airfoil shape. We can also slice a wing perpendicular to the leading edge or
to the quarter cord line or in fact perpendicular to any constant chord line. The resulting cross
section would be a different airfoil shape and would be referred to the airfoil perpendicular to x
chord position. In most cases we are interested in the stream-wise airfoil, also called the wing
section. If this section is then extended out to infinity, we can discuss the properties of the wing
section as if it were a two-dimensional (2-D) wing, and hence these are called 2-D airfoil
properties or 2-D wing section properties. We need the following three 2-D wing section
properties,
= 2-D lift curve slope
= 2-D zero-lift angle-of-attack
= 2-D pitch moment at zero lift =
This information is available for all NACA and some other standard airfoils. If we know some
of the geometry of an airfoil, we can estimate the 2-D lift curve slope in the following way:
2. 1) Estimate (measure is better) the trailing edge angle,
where is the thickness to chord ratio of the airfoil. Note that this estimate can be off by 300%!
2. Then calculate the theoretical 2-D lift-curve slope:
3. Modify the value in (2) to account for Reynolds number
Re = 106
Re = 107
4. Calculate the final 2-D lift-curve slope
(2)
where the term in parentheses is obtained from (3) and the last term from (2).
If no information is available, and the wing is thin, an approximate value of the 2-D lift
curve slope is
must be given or obtained by other means (eg. CFD)
3. Wing Geometry
The straight tapered
wing can be represented by a
few geometric properties that
describe the wing. We can
define some geometric
properties of the wing that
we will use throughout our
discussion. Again you should
be reminded that these are
characteristics of lifting
surfaces. Hence similar
properties can be assigned to
the horizontal tail but they
will have different values.
Taper Ratio,
Mean Geometric Chord, c
(3)
Aspect Ratio, AR
(4)
where b is the complete wing span, and S is the wing area.
Mean Aerodynamic Chord,
4. (5)
Half-Wing Lateral Mean Aerodynamic Chord Location,
(6)
Location of Leading Edge of Mean Aerodynamic Chord With Respect to Wing Apex
(Leading Edge of Root Chord),
(7)
Sweep of Any Fraction (%) Chord Line
In some of the estimation procedures to follow, the sweep angle of various chord lines are
required. We can define the nth fraction chord location, e.g. n = 1/4 for the quarter chord location.
Then if one fraction is represented by n, and another by m, then we can relate the sweep angle at
the nth location to the one at the mth location by:
(8)
Lifting Surface Lift Curve Slope
The objective of all this activity is to predict the lift curve slope of the wing (wb),
horizontal tail, and later the vertical tail surfaces. Hence the lift curve slope estimation presented
next must be applied several times, once for the wing using wing properties, once for the
horizontal tail using horizontal tail properties, and once for each other lifting surface using the
properties of that surface. The results given below are valid for all ranges of subsonic flight up
until the critical Mach number (that Mach number where somewhere on the aircraft the flow
becomes supersonic). The estimate is given by:
5. (9)
where x indicates the surface of interest, wb, ht, etc.
The value of is the ratio of the actual 2-D theoretical lift-curve slope with corrections
for compressibility (Mach number) effects divided by the flat plate theoretical lift curve slope
corrected for compressibility effects, . Consequently it becomes:
(10)
However, if we assume that the compressibility correction is the same for the actual as the
theoretical, e.g. , then we can write Eq. (10) in the following way:
(11)
where is the incompressible 2-D lift curve slope.
If is not given, and the wing is thin, then the theoretical 2-D lift curve slope can be
assume to be 2 and the value of k = 1!
Again, this estimate equation is valid for a wide range of aspect ratios and sweep angles,
and for the full range of subsonic Mach numbers up to the critical Mach number.
Supersonic Lift Curve Slope
For super sonic flight, the lift curve slope behaves differently. It can be given
approximately for thin wings as:
(12)
6. Wing Zero-Lift Angle-of-Attack,
Here it is assumed that we are given the zero-lift angle-of-attack of the wing section.
Then the zero-lift angle-of-attack for the complete wing can be determined for the following
cases:
1) Constant wing section (across the span), no sweep,
2) Constant wing section, sweep, wing section taken normal to x% chord line
(13)
where is the sweep angle of the x% chord line.
3) Constant wing section with twist. Reference all angles-of-attack to the root chord,
(think of ) for this calculation. Then,
where is the wing twist,( if linear it would be )
(14)
Wing
For this calculation we define the airfoil as that parallel to the free stream. It is assumed
we know the zero-lift pitch-moment for this 2D wing section. Then we can calculate the
complete wing zero-lift pitch-moment for the following cases:
1) Untwisted, constant section wings:
(15)
7. 2) Untwisted, varying wing section:
(16)
Wing Aerodynamic Center,
The location of the aerodynamic center on the mean aerodynamic chord can be best
obtained from charts developed for this purpose. These charts can present the results in different
ways. Typically the aerodynamic center position is referenced to the leading edge of the mean
aerodynamic chord, but sometimes its given with respect to the leading edge of the root chord or
wing apex. Etkin and Reid, Appendix C, Fig C.3 gives the aerodynamic center location for
various aspect ratios and wing taper ratios. Values that fall between the graphs or lines on the
graphs can be interpolated.
Downwash Parameter,
There are several ways to estimate this parameter. For straight tapered wings, a quasi-empirical,
analytic method has been devised that includes the effects of taper ratio, aspect ratio,
and horizontal tail vertical position. The method is presented in Etkin and Reid, Appendix B.5,
and is reproduced here with an example.
(17)
where: is a correction factor for the aspect ratio and is given by,
(18)
and is a correction factor for the taper ratio and is given by,
(19)
8. and is the correction factor for the location of the horizontal tail given by,
(20)
where: b is the wing span,
(modified tail length)
is the height above or below the plane formed by the root chord and the body
fixed y axis, or the distance measured in the plane of symmetry from the
root chord line to the horizontal tail aerodynamic center. (Not shown
correctly in Etkin and Reid, at least in my edition)
The result of evaluating Eq. (17) is the incompressible downwash parameter. It can be corrected
for compressibility effects using the standard compressibility correction (Prandtl-Glauert):
(21)
Again, these estimations are only good for subsonic flight speeds through the critical Mach
number.
Example: AR = 2.31, b = 36.5 ft = 0, =15.88 ft = 52.4 deg,
= 31.57 ft
This result is for the incompressible (low speed) case.