Turbulators on Helicopter Blades ?
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Turbulators on Helicopter Blades ?
Hi eng people.
do you know if a system similar to the "turbulators" on fixed wings has been tested on helicopter blades ?
if not what is your opinion on that ?
thank you.
Victor
do you know if a system similar to the "turbulators" on fixed wings has been tested on helicopter blades ?
if not what is your opinion on that ?
thank you.
Victor
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hi J.Cran
sorry for the word "turbulator".
I don't mean a vortex generator or a fence.
I'm trying to explain : on a wing, a line at the stall limit that "delimits" the stall region and catches the stall "bubble", or the early turbulence.
http://www.mh-aerotools.de/airfoils/turbulat.htm
this shows a pneumatic turbulator, but it can simply be a V groove.
From my little understanding,the are generally used to "drive" and delimit the stall zone , avoiding some vibrations.
i was thinking about that while i was thinking on a high-lift - airfoil , with a late but brutal stall.
has it been experimented ?
thank you
sorry for the word "turbulator".
I don't mean a vortex generator or a fence.
I'm trying to explain : on a wing, a line at the stall limit that "delimits" the stall region and catches the stall "bubble", or the early turbulence.
http://www.mh-aerotools.de/airfoils/turbulat.htm
this shows a pneumatic turbulator, but it can simply be a V groove.
From my little understanding,the are generally used to "drive" and delimit the stall zone , avoiding some vibrations.
i was thinking about that while i was thinking on a high-lift - airfoil , with a late but brutal stall.
has it been experimented ?
thank you
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Turbulators would probably cause mechnical stresses over and above the forces of an already highly stressed component (radially).
The aerodynamic side would also be complex because of the varying airspeeds and directions over the blade surfaces (each revolution of the rotor) when in forward flight.
The aerodynamic side would also be complex because of the varying airspeeds and directions over the blade surfaces (each revolution of the rotor) when in forward flight.
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sproket is right, the awful world of a helicopter blade is a far cry from that of an airplane wing in cruise. The properties of the airfoil are often selected as a compromise beyween the need for low drag in hover, high lift in high speed maneuvering, and low induced drag near stall. It is possible that the kinds of things that work on steady flow on an airplane have little positive effect when the blade goes from near transonic to near low speed stall in 150 milliseconds, and does that alternatively for hours at a time!
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Thanks,
i also agree with you sprocket.
i was comparing drag and lift polars of a VR7 and an ONERA OA213 and wondering what solutions had been tested to improve lift in the rev-flow zone or, at least, to control the stall.
The funny is that all numbers about the airfoils are given for a straight perpendicular airflow, but this "ideal" airflow just happens at the 90 deg azimut.. the rest of the time a blade being submitted to a mix of front and lateral flow. in fact i guess the air flow on an airfoil is more curved. or am i wrong?
If i'm correct how is it managed in the theories?
thanks.
i also agree with you sprocket.
i was comparing drag and lift polars of a VR7 and an ONERA OA213 and wondering what solutions had been tested to improve lift in the rev-flow zone or, at least, to control the stall.
The funny is that all numbers about the airfoils are given for a straight perpendicular airflow, but this "ideal" airflow just happens at the 90 deg azimut.. the rest of the time a blade being submitted to a mix of front and lateral flow. in fact i guess the air flow on an airfoil is more curved. or am i wrong?
If i'm correct how is it managed in the theories?
thanks.
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turbulator
zeeoo,
Turbulators are used mostly on very high performance sailplanes that have laminar flow airfoils.The purpose is for tripping the laminar bubble.
A strip of zig-zag tape is applied, usually aft of 50% chord.
A vortex generator is usually a piece of aluminum about 1"x .375" bonded every 6" or so.
Helicopter blades cannot maintain laminar flow, because for laminar flow an extremely smooth surface must be maintained.
The sandblasted finish of heli blades is far too rough. Even dust can destroy laminar flow.
slowrotor
Turbulators are used mostly on very high performance sailplanes that have laminar flow airfoils.The purpose is for tripping the laminar bubble.
A strip of zig-zag tape is applied, usually aft of 50% chord.
A vortex generator is usually a piece of aluminum about 1"x .375" bonded every 6" or so.
Helicopter blades cannot maintain laminar flow, because for laminar flow an extremely smooth surface must be maintained.
The sandblasted finish of heli blades is far too rough. Even dust can destroy laminar flow.
slowrotor
Turbulators when they blow tangential to vortex generators can sometimes trigger induced oscillations which perpetuate a resonance approximately 3 generations to the first harmonic. When this becomes fully established laminar fatigue envelopes are propgated at the genesis of the flow and allow boundary flow to disrupt.
It could reduce your Vne by up to 1 knot on a still wind day.
It could reduce your Vne by up to 1 knot on a still wind day.
Boundary Layer Transition Control
Zeeoo,
Turbulators are essentially boundary layer trips, which are used to induce laminar to turbulent transition further forward than it would have occured naturally. Boundary layer transition is a very complicated process for which there is no known general theory - even for simple geometries like a parallel flow flat plate case. A number of empirical relations are available but they have very limited applicability and cannot be applied willy-nilly.
The most important point to understand here is the way in which transition occurs on a 2D steady aerofoil at different Reynolds numbers. Let's take for example the venerable NACA0012 aerofoil, with which I have a lot of experience.
Firstly, for a medium size helicopter the tip Reynolds number in the hover is around 3,000,000. Reynolds number is simply the ratio of inertial and viscous forces in the flow. At these sorts of Reynolds numbers transition on isolated steady aerofoils generally occurs as the result of the Tollmien-Schlichting instability up until relatively high incidence when separation can become important. The NACA0012 does have a small separation bubble under these conditions, but it usually plays no part in the stall process. (See NASA TP1100) In addition for this Reynolds number and assuming a very low speed, say Mach 0.15, the transition on the upper surface will move from 0.21c to the leading edge between 3-12 degrees, much further forward than the low Reynolds number cases on the web site linked.
Now, at very low Reynolds number such as those found on low pressure compressor and turbine blades, glider wings and model aircraft, things change considerably. At very low Reynolds numbers transition would naturally occur a very long way aft. This results in the existence of a large region of laminar flow. As has rightly been pointed out; laminar flow is relatively unstable and cannot tolerate adverse pressure gradients or surface roughness. Therefore, a laminar separation often occurs. When this happens two events can follow. If transition occurs in the free shear layer, then the resulting turbulent boundary layer will be able to reattach creating a laminar separation bubble - as described in the web page zeeoo linked to. If transition, doesn't occur early enough then the flow will remain completely separated. This highlights the key difficulty in producing efficient aerodynamic shapes at low Reynolds numbers and is one of the key reasons why small scale gas turbine engines are inefficient and why model aircraft require such high power-to-weight ratios to make them work effectively.
In low Reynolds number flows, the turbulator provides a boundary layer trip that causes transition to occur much earlier than it would have done naturally. By triggering transition early, the laminar separation cannot occur and so the flow behaves much more like a high Reynolds number flow.
Clearly then, this technique can only add value in situation where the rotor is operating at very low Reynolds numbers, such as a Martian rotorcraft, and has been designed with an aerofoil that is completely inappropriate for that aerodynamic condition. At low Reynolds numbers, aerofoils must be carefully designed to control the large laminar region without triggering large separations. At high Reynolds numbers these problems seldom occur because transition occurs much closer to the leading edge (on the upper surface) and so transition occurs before laminar separation anyway.
In my opinion, turbulators are not appropriate for real helicopter/gyro plane flows, because the Reynolds numbers and Mach numbers will be too high for laminar separation be a serious problem - provided the aerofoil choices are sensible. At low Reynolds numbers they are a bit of a 'bodge-fix' and are a kind of admission of defeat in terms of designing a suitable aerofoil for the flight condition...we couldn't do it properly so we tripped the flow early to avoid the problem. Certainly not elegant engineering!
The other important point as Nick rightly pointed out is that there is a whole world of difference between the transition behaviour in steady and unsteady flows. I can tell you for certain that at typical helicopter Reynolds numbers and flight conditions that the transition behaviour in the leading edge region of the upper surface of common rotor aerofoil has a dramatic effect on the dynamic stall behaviour and the unsteady loads produced...but that's a heck of a lot more complicated!
So in answer to your question; has it been experimented with…yes, but not for the reasons you want. Will it help? No, not unless you are building a model gyrocopter! Will it help control stall? No, because stall is not a result of premature laminar separation on most helicopter aerofoils at realistic Reynolds numbers.
Hope this helps
CRAN
Turbulators are essentially boundary layer trips, which are used to induce laminar to turbulent transition further forward than it would have occured naturally. Boundary layer transition is a very complicated process for which there is no known general theory - even for simple geometries like a parallel flow flat plate case. A number of empirical relations are available but they have very limited applicability and cannot be applied willy-nilly.
The most important point to understand here is the way in which transition occurs on a 2D steady aerofoil at different Reynolds numbers. Let's take for example the venerable NACA0012 aerofoil, with which I have a lot of experience.
Firstly, for a medium size helicopter the tip Reynolds number in the hover is around 3,000,000. Reynolds number is simply the ratio of inertial and viscous forces in the flow. At these sorts of Reynolds numbers transition on isolated steady aerofoils generally occurs as the result of the Tollmien-Schlichting instability up until relatively high incidence when separation can become important. The NACA0012 does have a small separation bubble under these conditions, but it usually plays no part in the stall process. (See NASA TP1100) In addition for this Reynolds number and assuming a very low speed, say Mach 0.15, the transition on the upper surface will move from 0.21c to the leading edge between 3-12 degrees, much further forward than the low Reynolds number cases on the web site linked.
Now, at very low Reynolds number such as those found on low pressure compressor and turbine blades, glider wings and model aircraft, things change considerably. At very low Reynolds numbers transition would naturally occur a very long way aft. This results in the existence of a large region of laminar flow. As has rightly been pointed out; laminar flow is relatively unstable and cannot tolerate adverse pressure gradients or surface roughness. Therefore, a laminar separation often occurs. When this happens two events can follow. If transition occurs in the free shear layer, then the resulting turbulent boundary layer will be able to reattach creating a laminar separation bubble - as described in the web page zeeoo linked to. If transition, doesn't occur early enough then the flow will remain completely separated. This highlights the key difficulty in producing efficient aerodynamic shapes at low Reynolds numbers and is one of the key reasons why small scale gas turbine engines are inefficient and why model aircraft require such high power-to-weight ratios to make them work effectively.
In low Reynolds number flows, the turbulator provides a boundary layer trip that causes transition to occur much earlier than it would have done naturally. By triggering transition early, the laminar separation cannot occur and so the flow behaves much more like a high Reynolds number flow.
Clearly then, this technique can only add value in situation where the rotor is operating at very low Reynolds numbers, such as a Martian rotorcraft, and has been designed with an aerofoil that is completely inappropriate for that aerodynamic condition. At low Reynolds numbers, aerofoils must be carefully designed to control the large laminar region without triggering large separations. At high Reynolds numbers these problems seldom occur because transition occurs much closer to the leading edge (on the upper surface) and so transition occurs before laminar separation anyway.
In my opinion, turbulators are not appropriate for real helicopter/gyro plane flows, because the Reynolds numbers and Mach numbers will be too high for laminar separation be a serious problem - provided the aerofoil choices are sensible. At low Reynolds numbers they are a bit of a 'bodge-fix' and are a kind of admission of defeat in terms of designing a suitable aerofoil for the flight condition...we couldn't do it properly so we tripped the flow early to avoid the problem. Certainly not elegant engineering!
The other important point as Nick rightly pointed out is that there is a whole world of difference between the transition behaviour in steady and unsteady flows. I can tell you for certain that at typical helicopter Reynolds numbers and flight conditions that the transition behaviour in the leading edge region of the upper surface of common rotor aerofoil has a dramatic effect on the dynamic stall behaviour and the unsteady loads produced...but that's a heck of a lot more complicated!
So in answer to your question; has it been experimented with…yes, but not for the reasons you want. Will it help? No, not unless you are building a model gyrocopter! Will it help control stall? No, because stall is not a result of premature laminar separation on most helicopter aerofoils at realistic Reynolds numbers.
Hope this helps
CRAN
Last edited by CRAN; 25th Jan 2005 at 10:16.
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Thomas, Slow, J.Cran
great explanations.
I fully agree with you cran when you say it's a defeating engeneering
a kind of "driving the worse"..
I still having questions in the choice of a OA213 and a VR7 or NACA0012 , the first has, appearently a design made for higher AOA , similar to the VR4.
My questions would be what choice for thinner chords (lets say 16 cm). and what are the wayto reduce the reverse flow and gain efficiency.
I'm also wondering in the sanded surface on the upper surface... it sould be very slick... the sanded should be under, to slow the flow.. or is it a way to deal with a non-laminar but averaged up flow ?
Thanks to all, this really helps.
great explanations.
I fully agree with you cran when you say it's a defeating engeneering
a kind of "driving the worse"..I still having questions in the choice of a OA213 and a VR7 or NACA0012 , the first has, appearently a design made for higher AOA , similar to the VR4.
My questions would be what choice for thinner chords (lets say 16 cm). and what are the wayto reduce the reverse flow and gain efficiency.
I'm also wondering in the sanded surface on the upper surface... it sould be very slick... the sanded should be under, to slow the flow.. or is it a way to deal with a non-laminar but averaged up flow ?
Thanks to all, this really helps.
Zeeoo: you could consider postulating the disturbed flow under the nacelles and wingform.
This gives optimised non laminer flow without degradation of theta2.
try a reynolds number equivalent to the latest NACA series amendment?
This gives optimised non laminer flow without degradation of theta2.
try a reynolds number equivalent to the latest NACA series amendment?
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Thomas, thanks.
I misunderstand your first recommendation :
what disturbed flow ? you mean sucking it ?
what are the latest NASA series ? high lift ?
i heard about Re about 20M to 30M , is it right ?
thank you
I misunderstand your first recommendation :
what disturbed flow ? you mean sucking it ?
what are the latest NASA series ? high lift ?
i heard about Re about 20M to 30M , is it right ?
thank you
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Oh.. thanks Cran.
thanks for not having and ego so inflated by theta.
I'm not interested in studying the vaccum inside the skulls.
excuse my poor english.
BTW i woud be grateful to someone able to tell if there is an algorithm or approx. to calculate a Coeff of lift (Cz in my language), for a given airfoil.
Thanks.
thanks for not having and ego so inflated by theta
.I'm not interested in studying the vaccum inside the skulls.
excuse my poor english.
BTW i woud be grateful to someone able to tell if there is an algorithm or approx. to calculate a Coeff of lift (Cz in my language), for a given airfoil.
Thanks.
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zeeoo,
I looked for the very same information a few years back, but with little luck. There is an economical book called 'Theory of Wing Sections' by Abbott and Doenhoff, but it is from 1959. In addition, Prouty's main book gives quite a bit of information on the basic airfoils, such as the NACA 0012.
If you come across a newer or better source for this information, please advise.
Here is the coding, in Basic, for the lift for the NACA 0012. If you want, I can email you the whole boring module.
______________________________
Public Function coefficient_of_lift_NACA_0012(Alpha, AlphaL0, M, aa As Single) As Single
'NACA 0012
Dim AlphaL As Single 'Angle of attack where the lift coefficient
' first shows the effects of stall.
Dim K1 As Single 'Coefficient
Dim K2 As Single 'Coefficient
On Error GoTo coefficient_of_lift_NACA_0012_Err
'The basic simple equations.
'cl = aa * Alpha
'cl = 0.1 * Alpha 'Alpha (angle of attack) in degrees.
'cl = 6.0 * Alpha 'Alpha (angle of attack) in radians.
If (M < 0.725) Then ' Lift coefficients below 0.725M
AlphaL = 15 - 16 * M 'First effects of stall.
If (Alpha > AlphaL) Then 'If above stall.
K1 = 0.0233 + 0.342 * M ^ 7.15
K2 = 2.05 - 0.95 * M '## PROBLEM IF BELOW IS NEGATIVE & BELOW IS NOT A INTEGER
coefficient_of_lift_NACA_0012 = (aa * Alpha) - (K1 * ((Alpha - AlphaL) ^ K2))
Else 'If below stall.
coefficient_of_lift_NACA_0012 = aa * Alpha
End If
Else '(M < 0.725) Then 'Lift coefficients above 0.725M
K1 = 0.575 - 0.144 * (M - 0.725) ^ 0.44
K2 = 2.05 - 0.95 * M
coefficient_of_lift_NACA_0012 = ((0.677 - 0.744 * M) * Alpha) - (0.0575 - 0.144 * (M - 0.725) ^ 0.44) * ((Alpha - 3.4) ^ (2.05 - 0.95 * M))
End If
coefficient_of_lift_NACA_0012_Exit:
Exit Function
coefficient_of_lift_NACA_0012_Err:
MsgBox Err.Description
Resume coefficient_of_lift_NACA_0012_Exit
End Function:
I looked for the very same information a few years back, but with little luck. There is an economical book called 'Theory of Wing Sections' by Abbott and Doenhoff, but it is from 1959. In addition, Prouty's main book gives quite a bit of information on the basic airfoils, such as the NACA 0012.
If you come across a newer or better source for this information, please advise.
Here is the coding, in Basic, for the lift for the NACA 0012. If you want, I can email you the whole boring module.
______________________________
Public Function coefficient_of_lift_NACA_0012(Alpha, AlphaL0, M, aa As Single) As Single
'NACA 0012
Dim AlphaL As Single 'Angle of attack where the lift coefficient
' first shows the effects of stall.
Dim K1 As Single 'Coefficient
Dim K2 As Single 'Coefficient
On Error GoTo coefficient_of_lift_NACA_0012_Err
'The basic simple equations.
'cl = aa * Alpha
'cl = 0.1 * Alpha 'Alpha (angle of attack) in degrees.
'cl = 6.0 * Alpha 'Alpha (angle of attack) in radians.
If (M < 0.725) Then ' Lift coefficients below 0.725M
AlphaL = 15 - 16 * M 'First effects of stall.
If (Alpha > AlphaL) Then 'If above stall.
K1 = 0.0233 + 0.342 * M ^ 7.15
K2 = 2.05 - 0.95 * M '## PROBLEM IF BELOW IS NEGATIVE & BELOW IS NOT A INTEGER
coefficient_of_lift_NACA_0012 = (aa * Alpha) - (K1 * ((Alpha - AlphaL) ^ K2))
Else 'If below stall.
coefficient_of_lift_NACA_0012 = aa * Alpha
End If
Else '(M < 0.725) Then 'Lift coefficients above 0.725M
K1 = 0.575 - 0.144 * (M - 0.725) ^ 0.44
K2 = 2.05 - 0.95 * M
coefficient_of_lift_NACA_0012 = ((0.677 - 0.744 * M) * Alpha) - (0.0575 - 0.144 * (M - 0.725) ^ 0.44) * ((Alpha - 3.4) ^ (2.05 - 0.95 * M))
End If
coefficient_of_lift_NACA_0012_Exit:
Exit Function
coefficient_of_lift_NACA_0012_Err:
MsgBox Err.Description
Resume coefficient_of_lift_NACA_0012_Exit
End Function:
Zeeoo,
There is, but its not simple! To calculate the behaviour of the aerodynamics around any given shape for a viscous, compressible, unsteady flow you need to solve the Navier-Stokes equations. The implementation of a 2D Navier-Stokes solver for aerofoils, involves about 100-200 pages of code depending on its complexity. Once you have your code you will need to solve the equations on a suitable computational grid - you will need grid generation software to produce the grid also.
Once you have a code and a grid for the aerofoil, you can expect the calculation to take about half a day to accurately compute lift and drag, for a single design point on a desktop PC. (Less if you accept lower accuracy.)
If you accept lower accuracy you could also choose to solve reduced forms of the equations, which may, for example, limit you to attached incompressible flows. For this I suggest you do a search for XFOIL on the internet, it's free and should get you started! Plus it runs much, much quicker!
Incidentally, before you get too excited, the best practical useful methods available today (Reynolds Averaged Navier-Stokes Solvers) cannot accurately predict Clmax, i.e. the static stall point. Nor can they predict drag particularly well! It takes a very long time to learn how to tease useful engineering data out of these methods and if not applied very carefully will produce nonsense for results! Like I said...it's not easy!
Hope this helps
CRAN
PS: If I were designing a recreational gyrocopter I probably wouldn't be attempting to design new aerofoils - I would use established aerofoils for which there is lots of EXPERIMENTAL test data over a broad range of conditions. Perhaps the NACA0012...
Like Dave said 'The Theory of Wing Sections' by Abbott and Von Doenhoff is your best bet.
There is, but its not simple! To calculate the behaviour of the aerodynamics around any given shape for a viscous, compressible, unsteady flow you need to solve the Navier-Stokes equations. The implementation of a 2D Navier-Stokes solver for aerofoils, involves about 100-200 pages of code depending on its complexity. Once you have your code you will need to solve the equations on a suitable computational grid - you will need grid generation software to produce the grid also.
Once you have a code and a grid for the aerofoil, you can expect the calculation to take about half a day to accurately compute lift and drag, for a single design point on a desktop PC. (Less if you accept lower accuracy.)
If you accept lower accuracy you could also choose to solve reduced forms of the equations, which may, for example, limit you to attached incompressible flows. For this I suggest you do a search for XFOIL on the internet, it's free and should get you started! Plus it runs much, much quicker!
Incidentally, before you get too excited, the best practical useful methods available today (Reynolds Averaged Navier-Stokes Solvers) cannot accurately predict Clmax, i.e. the static stall point. Nor can they predict drag particularly well! It takes a very long time to learn how to tease useful engineering data out of these methods and if not applied very carefully will produce nonsense for results! Like I said...it's not easy!
Hope this helps
CRAN
PS: If I were designing a recreational gyrocopter I probably wouldn't be attempting to design new aerofoils - I would use established aerofoils for which there is lots of EXPERIMENTAL test data over a broad range of conditions. Perhaps the NACA0012...
Like Dave said 'The Theory of Wing Sections' by Abbott and Von Doenhoff is your best bet.
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Hi,
JCran, this helps and makes sense.
I don't try to resolve the Navier thing. as yous said, i just need a common, united database from whitch i could get the data to include in my routines. a kind of CLift (airfoil, AoA).
not much, as you said, i believe i must not go into a comlex simulation but re-use existing data but Whitch data.
XFoil and JAVAFoil are great but i can't get them interfaced with my code yet.
Dave, thanks for the code, interesting, but very particular to the naca0012, if i understand it correctly. i asked the same on rotary and got an approx solution for a 3D wing , i think it is for a very approx aprox :
---------------------------------- from Gabriel Hugh Elkaim
- rotary
Cl = 2*pi*AR/(AR+2)*alpha
where alpha is the angle of attack from the zero lift line in RADIANS
AR is the aspect ratio (span/chord for a rectangular wing)
____________________________________
using the AR to calculatet hat is a kind of simple , how "un-acurate" can it be ?
for your understanding : i try to code a simulator that will compute the lift along a blade, with a transition between different airfoils and different twist values. Then i will compute the lift on the disk at different angles and airfoil combinations.
Since I'm not a math man, i slice the blade in N small elements to plot the lift/drag values.
QUESTIONS :
I was wondering if i can interlopate an average CLift betwen 2 airfoils CLifts at the same AoA.
How to consider the airflow on the 0 degree azimut (blade at the front), the blade being submitted to an angular speed + a frontal speed, the air flow passes through a distorted airfoil. could i introduce a kind of coefficient of distortion ? or can i consider the lift is the same.
Thanks for your quality help.
---------------------------------- from Geneweber - rotary
The Department of Aerospace Engineering at the University of Illinois has a large airfoil coordinate database. http://www.aae.uiuc.edu/m-selig/ads.html I ran some of these through both JavaFoil http://www.mh-aerotools.de/airfoils/javafoil.htm and Xfoil http://raphael.mit.edu/xfoil/ and was able to generate lift and drag plots. It was interesting. Hope this helps.
JCran, this helps and makes sense.
I don't try to resolve the Navier thing. as yous said, i just need a common, united database from whitch i could get the data to include in my routines. a kind of CLift (airfoil, AoA).
not much, as you said, i believe i must not go into a comlex simulation but re-use existing data but Whitch data.
XFoil and JAVAFoil are great but i can't get them interfaced with my code yet.
Dave, thanks for the code, interesting, but very particular to the naca0012, if i understand it correctly. i asked the same on rotary and got an approx solution for a 3D wing , i think it is for a very approx aprox :
---------------------------------- from Gabriel Hugh Elkaim
- rotary
Cl = 2*pi*AR/(AR+2)*alpha
where alpha is the angle of attack from the zero lift line in RADIANS
AR is the aspect ratio (span/chord for a rectangular wing)
____________________________________
using the AR to calculatet hat is a kind of simple , how "un-acurate" can it be ?
for your understanding : i try to code a simulator that will compute the lift along a blade, with a transition between different airfoils and different twist values. Then i will compute the lift on the disk at different angles and airfoil combinations.
Since I'm not a math man, i slice the blade in N small elements to plot the lift/drag values.
QUESTIONS :
I was wondering if i can interlopate an average CLift betwen 2 airfoils CLifts at the same AoA.
How to consider the airflow on the 0 degree azimut (blade at the front), the blade being submitted to an angular speed + a frontal speed, the air flow passes through a distorted airfoil. could i introduce a kind of coefficient of distortion ? or can i consider the lift is the same.
Thanks for your quality help.
---------------------------------- from Geneweber - rotary
The Department of Aerospace Engineering at the University of Illinois has a large airfoil coordinate database. http://www.aae.uiuc.edu/m-selig/ads.html I ran some of these through both JavaFoil http://www.mh-aerotools.de/airfoils/javafoil.htm and Xfoil http://raphael.mit.edu/xfoil/ and was able to generate lift and drag plots. It was interesting. Hope this helps.
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zeeoo,
" i try to code a simulator that will compute the lift along a blade, with a transition between different airfoils and different twist values. Then i will compute the lift on the disk at different angles and airfoil combinations.
Since I'm not a math man, i slice the blade in N small elements to plot the lift/drag values."
What you are looking for is in Prouty's book 'Helicopter Performance, Stability and Control". He provides 21 steps, and their algorithms, to calculate the lift of rotors and their blades, in hover. The calculations are based on 'Combined Momentum and Blade Element Theory with Empirical Corrections'.
The inputs are;
Rotor geometry; - number of blades, radius, chord, twist, cutout and airfoil data.
Test conditions;- tip speed, atmospheric density, speed of sound.
I have coded it in Basic in Microsoft's Access database. The previous post was the NACA 0012 lift calculations Module from this program.
The problem was that I could only find valid algorithms for the NACA0012 and the 8-H-12 airfoils. The algorithms for other airfoils could only be guesstimated.
If you can find the algorithims for other airfoils, I'll trade the coding and forms for them.
_______
These pages will have some information, which you know of, and perhaps some that is new to you.
OTHER: Aerodynamic - Airfoil
OTHER: Aerodynamic - Blade Profiles
Dave
" i try to code a simulator that will compute the lift along a blade, with a transition between different airfoils and different twist values. Then i will compute the lift on the disk at different angles and airfoil combinations.
Since I'm not a math man, i slice the blade in N small elements to plot the lift/drag values."
What you are looking for is in Prouty's book 'Helicopter Performance, Stability and Control". He provides 21 steps, and their algorithms, to calculate the lift of rotors and their blades, in hover. The calculations are based on 'Combined Momentum and Blade Element Theory with Empirical Corrections'.
The inputs are;
Rotor geometry; - number of blades, radius, chord, twist, cutout and airfoil data.
Test conditions;- tip speed, atmospheric density, speed of sound.
I have coded it in Basic in Microsoft's Access database. The previous post was the NACA 0012 lift calculations Module from this program.
The problem was that I could only find valid algorithms for the NACA0012 and the 8-H-12 airfoils. The algorithms for other airfoils could only be guesstimated.
If you can find the algorithims for other airfoils, I'll trade the coding and forms for them.
_______
These pages will have some information, which you know of, and perhaps some that is new to you.
OTHER: Aerodynamic - Airfoil
OTHER: Aerodynamic - Blade Profiles
Dave
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Hi Dave,
for airfoil data, i plan to plot myself some values, coming from an array, like you did on the vR7 data on your site (yes i read it sometimes).
I dont need an universal alrorithm, only some airfoils like VR7b, VR9, VR5, OA213(similar to VR12),B29-root , so, i think i will "hard" code the data, based on Bezier splines, i will "build" the curve my self trying to be as close as the original, doing so, i can have an interpolated (quite precise) data.
do you think i can interpolate between the CL of 2 given airfoils at the same AoA ? for example from 0.50 to 0.58, giving 0.54 ?
What i want tyo do is calculate the lift not in hover, but with a disk pitch and a relative wind (speed).
then , if i get it working, i will compute the vectors for each "slice" of the blade and will have an approx of the autorotation capabilities.
I can easily have a 3D interface, could be interesting.
then why not the calculation of the moments, centrifugal loads, and maybe compute the coning and the sissors load at the root.
and why not imput a kind of hinge elasticity , and why not a pitch imput... and why not...
i can get busy thill the summer
what do you think ?
I will never be 100% acurate (no method is), so i will deal with some approx.
BTW do you know if Prouty's book has been released in french?
thanks
for airfoil data, i plan to plot myself some values, coming from an array, like you did on the vR7 data on your site (yes i read it sometimes
).I dont need an universal alrorithm, only some airfoils like VR7b, VR9, VR5, OA213(similar to VR12),B29-root , so, i think i will "hard" code the data, based on Bezier splines, i will "build" the curve my self trying to be as close as the original, doing so, i can have an interpolated (quite precise) data.
do you think i can interpolate between the CL of 2 given airfoils at the same AoA ? for example from 0.50 to 0.58, giving 0.54 ?
What i want tyo do is calculate the lift not in hover, but with a disk pitch and a relative wind (speed).
then , if i get it working, i will compute the vectors for each "slice" of the blade and will have an approx of the autorotation capabilities.
I can easily have a 3D interface, could be interesting.
then why not the calculation of the moments, centrifugal loads, and maybe compute the coning and the sissors load at the root.
and why not imput a kind of hinge elasticity , and why not a pitch imput... and why not...
i can get busy thill the summer
what do you think ?
I will never be 100% acurate (no method is), so i will deal with some approx.
BTW do you know if Prouty's book has been released in french?
thanks
Last edited by zeeoo; 26th Jan 2005 at 19:48.
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hi engies,
i found this , it is not a turbulator but looks like a similar device, could anyone explein please ? Thanks
NUMERICAL STUDY OF MULTI-ELEMENT AIRFOIL
AERODYNAMICS
CLIN M. WANG Georgia Inst. of Tech., GA, US and CHEE TUNG In
Developments in theoretical and applied mechanics; Southeastern Conference
on Theoretical and Applied Mechanics, 16th, Nashville, TN, April
12-14, 1992. A95-93700 Tullahoma, TN The Univ. of Tennessee Space
Inst. (SECTAM, Vol. 16) 1992 p. III.II.32-III.II.39
(ISBN 1-879921-01-4) Copyright
Unsteady flowfields around oscillating Boeing VR7 airfoil with and
without a leading-edge slat were numerically investigated by a novel zonal method using a conformal mapping technique. Numerical aero-dynamic hysteresis loops show that the leading-edge slat prevents the airfoil dynamic stall at reduced frequency of 0.15, Reynolds number of 1 million
Question : what is this "slat" ? it's shape ?
i found this , it is not a turbulator but looks like a similar device, could anyone explein please ? Thanks
NUMERICAL STUDY OF MULTI-ELEMENT AIRFOIL
AERODYNAMICS
CLIN M. WANG Georgia Inst. of Tech., GA, US and CHEE TUNG In
Developments in theoretical and applied mechanics; Southeastern Conference
on Theoretical and Applied Mechanics, 16th, Nashville, TN, April
12-14, 1992. A95-93700 Tullahoma, TN The Univ. of Tennessee Space
Inst. (SECTAM, Vol. 16) 1992 p. III.II.32-III.II.39
(ISBN 1-879921-01-4) Copyright
Unsteady flowfields around oscillating Boeing VR7 airfoil with and
without a leading-edge slat were numerically investigated by a novel zonal method using a conformal mapping technique. Numerical aero-dynamic hysteresis loops show that the leading-edge slat prevents the airfoil dynamic stall at reduced frequency of 0.15, Reynolds number of 1 million
Question : what is this "slat" ? it's shape ?