Its not rocket science...or is it?
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Spanner Turner has it mostly right - but wait, there's more! 
The compressor (& fan) rotor blades impart the energy to the air, but it exists in the form of a helical or vortex flow. In other words it's not flowing straight aft, but is rotating in the direction the rotor throws it.
SO - behind each rotor stage there is a stator stage. The stator vanes (think of venetian blinds) straighten the airflow to a straight-aft direction. In so doing, they recover some static pressure (they behave like a diffuser), and there is a fair amount of additional thrust recovered (from the lift vector of the stator airfoil).
The rotor won't work right without the stator.
The compressor (& fan) rotor blades impart the energy to the air, but it exists in the form of a helical or vortex flow. In other words it's not flowing straight aft, but is rotating in the direction the rotor throws it.
SO - behind each rotor stage there is a stator stage. The stator vanes (think of venetian blinds) straighten the airflow to a straight-aft direction. In so doing, they recover some static pressure (they behave like a diffuser), and there is a fair amount of additional thrust recovered (from the lift vector of the stator airfoil).
The rotor won't work right without the stator.
Last edited by barit1; 8th Feb 2008 at 14:25.
Thanks everyone for the quick reaction (
). I'm nearly there, but I need some more information.
A special thank you to Capt Pit Bull who made my day:
PBL, I had discovered the NASA site after I posted and it is very helpful.
On the first page, the EPR formula is quoted:
I can see that the compressor pressure ratio is greater than 1.0 and therefore exerts a forward force. The compressor forces are transmitted through thrust bearings and the stator blade tip forces directly to the engine casing. What about the stator root forces? Through the thrust bearing support structure?
I can also see that the turbine will exert a backward force on the engine, again through thrust bearings (and stators?).
What about the burner cans and the nozzle? The NASA Java model default settings show no thrust for either of these (PR = 1.0). I thought the nozzle in particular was carefully designed to provide a thrust increment. If so, what does this thrust react against?
). I'm nearly there, but I need some more information.A special thank you to Capt Pit Bull who made my day:
Rocket engineering on the other hand is quite tricky
On the first page, the EPR formula is quoted:
EPR = pt8 / pt2 = (pt3 / pt2) * (pt4 / pt3) * (pt5 / pt4) * (pt8 / pt5)
EPR = compressor pressure ratio * burner pressure ratio * turbine pressure ratio * nozzle pressure ratio
EPR = compressor pressure ratio * burner pressure ratio * turbine pressure ratio * nozzle pressure ratio
I can also see that the turbine will exert a backward force on the engine, again through thrust bearings (and stators?).
What about the burner cans and the nozzle? The NASA Java model default settings show no thrust for either of these (PR = 1.0). I thought the nozzle in particular was carefully designed to provide a thrust increment. If so, what does this thrust react against?
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India Four Two, it's not so simple. The NASA diagram speaks of PRESSURES, but what we need to know is FORCES, and a force is = Pressure times Area.
And every pressure is acting on a different area. Doing the sums is quite a bookkeeping exercise, I've done it a time or two, and in the preliminary design phase the uncertainty of each element in the summation is so great that you're never quite sure which direction the thrust bearing is loaded - fore or aft!
(Think of a tug of war between two teams of elephants...)
And every pressure is acting on a different area. Doing the sums is quite a bookkeeping exercise, I've done it a time or two, and in the preliminary design phase the uncertainty of each element in the summation is so great that you're never quite sure which direction the thrust bearing is loaded - fore or aft!
(Think of a tug of war between two teams of elephants...)
And I assume in the end, the team of elephants at the pointy-end wins.
Doh 
I knew that! My only excuse is it's very late where I am.
I'm still interested in the nozzle thrust and what it reacts upon.
Pressure times Area
I knew that! My only excuse is it's very late where I am.I'm still interested in the nozzle thrust and what it reacts upon.
Oh dear we are about to get away from science and into engineering with this now.
Following along with the posts above, the pushies are both forward and aft, the reaction to these pushes initial are against engine structures which react by deforming ever so slightly before fighting it out as to which way they are going to push on the engine mounts (hopefully in the forward direction if you want to go forward).
One of the biggest forward pushes is the blades to the rotor disk to the shaft and thence to the ball bearing or thrust bearing. But leaving that single bearing to take all that load at the highest thurst condition while at the same time it might even end up with a reverse load in the other direction at idle is not agood thing for the bearing. It would rattle around at the low load condition and wear out quite rapidly. So good engineering steps in and decides that one must always have a positive moderate load on that bearing regardless of the power or engine thrust condition, so the engineers move some of the presures arround inside the machine (bleeds that go from the back of the compressor to anywhere in the innards of the engine that they see fit) and presto they can actually reverse the thrust load on the bearing to push aft in all operating conditions. So with this being the case the thrust load that gets to the mounts has got to be recovered from the momentum of the air through another structure other than the bearing.
So the only time you can really predict which way the rotor is pushing is if you break the shaft between the bearings.
Following along with the posts above, the pushies are both forward and aft, the reaction to these pushes initial are against engine structures which react by deforming ever so slightly before fighting it out as to which way they are going to push on the engine mounts (hopefully in the forward direction if you want to go forward).
One of the biggest forward pushes is the blades to the rotor disk to the shaft and thence to the ball bearing or thrust bearing. But leaving that single bearing to take all that load at the highest thurst condition while at the same time it might even end up with a reverse load in the other direction at idle is not agood thing for the bearing. It would rattle around at the low load condition and wear out quite rapidly. So good engineering steps in and decides that one must always have a positive moderate load on that bearing regardless of the power or engine thrust condition, so the engineers move some of the presures arround inside the machine (bleeds that go from the back of the compressor to anywhere in the innards of the engine that they see fit) and presto they can actually reverse the thrust load on the bearing to push aft in all operating conditions. So with this being the case the thrust load that gets to the mounts has got to be recovered from the momentum of the air through another structure other than the bearing.
So the only time you can really predict which way the rotor is pushing is if you break the shaft between the bearings.
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Many text books give diagrams to illustrate where the thrust is produced.
In their book "The Jet Engine" Rolls Royce give the follwing figures for "a typical single-spool axial flow engine".
Forward gas loads (producing forward thrust)
Comressor 19049 lbs
Diffuser (at the rear of the compressor) 2186 lbs
Combustion chambers 34182 lbs
Exhaust unit (imedaitaley aft of the turbines) 2419 lbs
TOTAL FORWARD GAS LOADS 57836 lbs
Rearward facing gas loads (producing drag)
Turbines 41091 lbs
Propelling Nozzle 5587 labs
TOTAL REARWARD GAS LOADS 46678 lbs
Total thrust = 57836 - 46678 = 11158 lbs
So in this type of engine the majority of the thrust comes form the air accelerating rearwards as it expands in the combustion chambers.
The next highest thrust producer is the compressor.
If we look at what was added to create the modern high thrust turbofans it was an enormous fan at the front. This should tell us that in these engines most of the thrust is created by the fan. This is certainly true at low altitude and low airspeed.
But as TAS and altitude increase, the contribution made by the fans decreases rapidly. In high altitude high speed cruise flight most of the thrust comes from the engine core.
In their book "The Jet Engine" Rolls Royce give the follwing figures for "a typical single-spool axial flow engine".
Forward gas loads (producing forward thrust)
Comressor 19049 lbs
Diffuser (at the rear of the compressor) 2186 lbs
Combustion chambers 34182 lbs
Exhaust unit (imedaitaley aft of the turbines) 2419 lbs
TOTAL FORWARD GAS LOADS 57836 lbs
Rearward facing gas loads (producing drag)
Turbines 41091 lbs
Propelling Nozzle 5587 labs
TOTAL REARWARD GAS LOADS 46678 lbs
Total thrust = 57836 - 46678 = 11158 lbs
So in this type of engine the majority of the thrust comes form the air accelerating rearwards as it expands in the combustion chambers.
The next highest thrust producer is the compressor.
If we look at what was added to create the modern high thrust turbofans it was an enormous fan at the front. This should tell us that in these engines most of the thrust is created by the fan. This is certainly true at low altitude and low airspeed.
But as TAS and altitude increase, the contribution made by the fans decreases rapidly. In high altitude high speed cruise flight most of the thrust comes from the engine core.
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Figuring out where thust comes from gets to be real fun when looking at supersonic flight. It goes a long way from "Throw lots of hot air out the back and go the other way"
The concordes engines produced around 75% of their thrust at M2.2 in the diffusor section and only around 8% came from the actual engine. Another 29% thrust came from the nozzles, whereas the supersonic shocks at the intakes prosuced around 12% drag.
In a way the aircraft sucked its way forward.
OORW
The concordes engines produced around 75% of their thrust at M2.2 in the diffusor section and only around 8% came from the actual engine. Another 29% thrust came from the nozzles, whereas the supersonic shocks at the intakes prosuced around 12% drag.
In a way the aircraft sucked its way forward.
OORW
This is what I love about PPRuNe. I probably could have found the answers online if I had searched long enough or if I had access to a decent library - not possible here in SGN.
However, by posting my question, I've got some interesting answers and a fact previously unknown to me.
At Mach 2.2, "Concorde sucks"!
Keith, I'm even more confused about the nozzle now. You are saying everything aft of the turbine is a net drag force? What's the purpose of the careful design of the nozzle shape? To reduce the drag to a minimum?
lomapaseo, are the thrust bearing bleeds on all the time, or are they turned off when the engine is under power?
However, by posting my question, I've got some interesting answers and a fact previously unknown to me.
At Mach 2.2, "Concorde sucks"!
Keith, I'm even more confused about the nozzle now. You are saying everything aft of the turbine is a net drag force? What's the purpose of the careful design of the nozzle shape? To reduce the drag to a minimum?
lomapaseo, are the thrust bearing bleeds on all the time, or are they turned off when the engine is under power?
Last edited by India Four Two; 10th Feb 2008 at 07:26.
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No, I did not say that "everything aft of the turbine is a net drag force". If you look carefully you will see that the RR figures include the statement that, "Exhaust unit (imediately aft of the turbines) 2419 lbs forward force". This is produced by the fact that pressure immediately behind the turbines is greater than ambient.
I we look only at these figures we may well ask, "so why don't we just get rid of the propelling nozzle?" Part of the answer is that if we removed the jet pipe and convergent propelling nozzle, the pressure behind the turbine would be lost as the air axpanded in all directions. This would cause us to lose the thrust that was being produced by the exhaust unit.
The jet pipe and propelling nozzle direct this expansion aft, and in this way it uses the excess pressure in the jet pipe to accelerate the air rearwards. This rearward acceleration produces thrust. In the RR figures this appears as the thrust on the exhaust unit.
If we look carefully at these nozzles in supersonic flight, the gas pressure on the convergent section tends to push the nozle aft, but that on the divergent section tends to push it forward.
But the excess pressure in the jet pipe is acting on the forward face of the propelling, nozzle so if the nozzle were to be suddenly released it would fly out of the back of the engine. This shows that the force on the propelling nozzle is acting reawards as stated in the RR figures.
But that is only part of the answer. The greatest benefit of the propelling nozzle becomes evident only as the speed of the aeroplane increases.
According to Newton's Second Law Thrust = Mass x Acceleration.
This means that to create thrust, the velocity of the air flowing out of the exhaust must be greater than that flowing in at the air inlet. In effect this means that the amount of thrust produced is proportional to the difference between the exhaust velocity and the TAS of the aeroplane. This in turn means that when TAS is equal to exhaust velocity there will be no thrust.
By accelerating the exhaust gas rearwards, the convergent propelling nozzle increases the thrust and also increases the TAS at which the thrust falls to zero. This permits us to reach higher airspeeds.
But convergent propelling nozzles cause acceleration only in subsonic airflows. So for very high supersonic speeds we need to replace the convergent propelling nozzle with one that is convergent-divergent. In these, the convergent part accelerates the air up to sonic speed, then the divergent part continues the acceleration at supersonic speeds.
The statement that "Concorde intakes suck the aircraft forward in supersonic flight" is incorrect.
For this idea to be true it would require the pressure inside the intake to be lower than ambient pressure. This would cause air to be sucked into the intake. But in high speed flight the air is rammed into forward facing intakes and the pressure inside the intake is greater than ambient.
A supersonic air intake is made up of a convegent duct, followed by a divergent duct. As supersonic air is rammed into the convergent section, its velocity decreases and its pressure increases. This produces an aft facing force (drag) on the convergent surfaces.
The convergent section is followed by a divergent section leading to the engine. By the time the air reaches the divergent section, its velocity is subsonic. So the air then continues to decelerate subsonically through the divergent section. This deceleration increases the pressure further.
Inside the divergent section the high pressure tends to push the divergent surfaces forward. It is this which produces the extra thrust. The thrust produced by the complete intake is equal to the forward force on the divergent surfaces minus the rearward force acting on the convergent surface. But there is no "sucking" involved in this process.
It might however be argued that at the start of the take-off run all jet engines suck the aircraft forward to some extent, because pressure in the inlet really is lower than ambient.
I we look only at these figures we may well ask, "so why don't we just get rid of the propelling nozzle?" Part of the answer is that if we removed the jet pipe and convergent propelling nozzle, the pressure behind the turbine would be lost as the air axpanded in all directions. This would cause us to lose the thrust that was being produced by the exhaust unit.
The jet pipe and propelling nozzle direct this expansion aft, and in this way it uses the excess pressure in the jet pipe to accelerate the air rearwards. This rearward acceleration produces thrust. In the RR figures this appears as the thrust on the exhaust unit.
If we look carefully at these nozzles in supersonic flight, the gas pressure on the convergent section tends to push the nozle aft, but that on the divergent section tends to push it forward.
But the excess pressure in the jet pipe is acting on the forward face of the propelling, nozzle so if the nozzle were to be suddenly released it would fly out of the back of the engine. This shows that the force on the propelling nozzle is acting reawards as stated in the RR figures.
But that is only part of the answer. The greatest benefit of the propelling nozzle becomes evident only as the speed of the aeroplane increases.
According to Newton's Second Law Thrust = Mass x Acceleration.
This means that to create thrust, the velocity of the air flowing out of the exhaust must be greater than that flowing in at the air inlet. In effect this means that the amount of thrust produced is proportional to the difference between the exhaust velocity and the TAS of the aeroplane. This in turn means that when TAS is equal to exhaust velocity there will be no thrust.
By accelerating the exhaust gas rearwards, the convergent propelling nozzle increases the thrust and also increases the TAS at which the thrust falls to zero. This permits us to reach higher airspeeds.
But convergent propelling nozzles cause acceleration only in subsonic airflows. So for very high supersonic speeds we need to replace the convergent propelling nozzle with one that is convergent-divergent. In these, the convergent part accelerates the air up to sonic speed, then the divergent part continues the acceleration at supersonic speeds.
The statement that "Concorde intakes suck the aircraft forward in supersonic flight" is incorrect.
For this idea to be true it would require the pressure inside the intake to be lower than ambient pressure. This would cause air to be sucked into the intake. But in high speed flight the air is rammed into forward facing intakes and the pressure inside the intake is greater than ambient.
A supersonic air intake is made up of a convegent duct, followed by a divergent duct. As supersonic air is rammed into the convergent section, its velocity decreases and its pressure increases. This produces an aft facing force (drag) on the convergent surfaces.
The convergent section is followed by a divergent section leading to the engine. By the time the air reaches the divergent section, its velocity is subsonic. So the air then continues to decelerate subsonically through the divergent section. This deceleration increases the pressure further.
Inside the divergent section the high pressure tends to push the divergent surfaces forward. It is this which produces the extra thrust. The thrust produced by the complete intake is equal to the forward force on the divergent surfaces minus the rearward force acting on the convergent surface. But there is no "sucking" involved in this process.
It might however be argued that at the start of the take-off run all jet engines suck the aircraft forward to some extent, because pressure in the inlet really is lower than ambient.
Last edited by Keith.Williams.; 10th Feb 2008 at 09:23.
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This was also one of my fixed dilemma... 
Is there someone that can explain why a High by pass turbofan is not the same as a propeller???
When we study propellers we know that they are affected by the efficiency of propulsion related to the a/c speed. Jet engine instead aren't.
Why the fan stage is not considered to be the same as a propeller.... the only difference looks to be the number of blades!!!! hence it should behave as a propeller!!!!
The turbofan should have a great dependency on the a/c speed too... instead it doesn't.... why?
Bye
Is there someone that can explain why a High by pass turbofan is not the same as a propeller???
When we study propellers we know that they are affected by the efficiency of propulsion related to the a/c speed. Jet engine instead aren't.
Why the fan stage is not considered to be the same as a propeller.... the only difference looks to be the number of blades!!!! hence it should behave as a propeller!!!!
The turbofan should have a great dependency on the a/c speed too... instead it doesn't.... why?
Bye
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By Newton's second law we have Thrust = Mass x Acceleration.
If we take a mass of air and exert a force on it such that we accelerate in rearwards, we will get forward thrust.
The amount of thrust produced is equal to the mass of air multiplied by the acceleration that we give to it.
We can take a very large mass and give it a small acceleration as in a propeller system. Or we can take a very small mass and give it a very large acceleration as we do in a pure jet (no by-pass). Or we can take a moderate mass and give it a moderate acceleration, as we do in a turbofan. In each case provided the mass multilied by the acceleration is the same, then the thrust produced will be the same.
But in real propulsion systems it is all a continuous process, so we can modify the equation to read
Thrust = Mass Aiflow (in Kg/sec) x (Exhaust speed - TAS (in m/sec))
Looking at this equation and assuming that the mass flow rate is constant, as we accelerate the aircraft and our TAS gets closer to the exhaust speed, the thrust that we are producing decreases. When we are flying at our own exhaust speed, we have no thrust.
For propeller systems the exhaust speed (propwash) is not very high, so we do not get to go very fast before losing most of the thrust.
In non-by-pass turbojet the exhaust velocity is extremely high, so we can go much faster. And as our TAS increases, a second factor comes in to play.
The ramming of the air into the intake increases its density. This in turn increases the mass airflow through the engine. The rate of increase is proportional to dynamic pressure which is exponential. This increase in mass flow more than compensates for the reduced acceleration. So as a pure jet aircraft accelerates, its thrust decreases up to about 250 knots TAS then starts to increase again as increasing ram effect becomes dominant.
For the purposes of Principles of Flight studies we usually pretend that it is constant. But for subjects such as take-off performance we must take into account the fact that take-off thrust actually decreases during the take-off run.
The high by-pass turbofan is somewhere between the prop and the pure jet.
Because it has an air intake it benefits to some extent from ram effect. But the velocity of the air coming out of the by-pass duct is much lower than that of a pure jet. So the benefit of ram efect is not so great.
The overall effect is that as TAS increases, the thrust produced by the by-pass flow decreases, but at a slower rate than that for a propeller. The thrust from the hot gas stream responds pretty much like that of a pure jet. Adding these two effects together we find that thrust in a high by-pass torbojet reduces with inceasing TAS but is still usable up to higher airspeeds than those which can reached with a propeller.
If we take a mass of air and exert a force on it such that we accelerate in rearwards, we will get forward thrust.
The amount of thrust produced is equal to the mass of air multiplied by the acceleration that we give to it.
We can take a very large mass and give it a small acceleration as in a propeller system. Or we can take a very small mass and give it a very large acceleration as we do in a pure jet (no by-pass). Or we can take a moderate mass and give it a moderate acceleration, as we do in a turbofan. In each case provided the mass multilied by the acceleration is the same, then the thrust produced will be the same.
But in real propulsion systems it is all a continuous process, so we can modify the equation to read
Thrust = Mass Aiflow (in Kg/sec) x (Exhaust speed - TAS (in m/sec))
Looking at this equation and assuming that the mass flow rate is constant, as we accelerate the aircraft and our TAS gets closer to the exhaust speed, the thrust that we are producing decreases. When we are flying at our own exhaust speed, we have no thrust.
For propeller systems the exhaust speed (propwash) is not very high, so we do not get to go very fast before losing most of the thrust.
In non-by-pass turbojet the exhaust velocity is extremely high, so we can go much faster. And as our TAS increases, a second factor comes in to play.
The ramming of the air into the intake increases its density. This in turn increases the mass airflow through the engine. The rate of increase is proportional to dynamic pressure which is exponential. This increase in mass flow more than compensates for the reduced acceleration. So as a pure jet aircraft accelerates, its thrust decreases up to about 250 knots TAS then starts to increase again as increasing ram effect becomes dominant.
For the purposes of Principles of Flight studies we usually pretend that it is constant. But for subjects such as take-off performance we must take into account the fact that take-off thrust actually decreases during the take-off run.
The high by-pass turbofan is somewhere between the prop and the pure jet.
Because it has an air intake it benefits to some extent from ram effect. But the velocity of the air coming out of the by-pass duct is much lower than that of a pure jet. So the benefit of ram efect is not so great.
The overall effect is that as TAS increases, the thrust produced by the by-pass flow decreases, but at a slower rate than that for a propeller. The thrust from the hot gas stream responds pretty much like that of a pure jet. Adding these two effects together we find that thrust in a high by-pass torbojet reduces with inceasing TAS but is still usable up to higher airspeeds than those which can reached with a propeller.
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The concordes engines produced around 75% of their thrust at M2.2 in the diffusor section and only around 8% came from the actual engine. Another 29% thrust came from the nozzles, whereas the supersonic shocks at the intakes prosuced around 12% drag.
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soundlover - Good question. See prior thread
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Originally Posted by OutOfRunWay
The Concorde's engines produced around 75% of their thrust at M2.2 in the diffusor section and only around 8% came from the actual engine. Another 29% thrust came from the nozzles, whereas the supersonic shocks at the intakes produced around 12% drag.
On another forum we wondered about the 12% drag figure. The 75/8/29 % figures would have been derived from the pressure ratios x areas, and therefore would already have taken internal drag into acount. We reckoned the 12% was the external drag of the nacelle structure.
uniuniunium,
One of my SR-71 books quotes even more extreme figures, and also states that under certain flight conditions (throttling back at Mach 3, IIRC) the engine produced no thrust as such at all, and actually moved back on the engine mountings....
Thanks to the contributors !
There IS a difference between having gotten your own head satisfactorily around the various notions and being able to describe them clearly and concisely to somebody else.
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When we study propellers we know that they are affected by the efficiency of propulsion related to the a/c speed. Jet engine instead aren't.
Why the fan stage is not considered to be the same as a propeller.... the only difference looks to be the number of blades!!!! hence it should behave as a propeller!!!!
The turbofan should have a great dependency on the a/c speed too... instead it doesn't.... why?
Why the fan stage is not considered to be the same as a propeller.... the only difference looks to be the number of blades!!!! hence it should behave as a propeller!!!!
The turbofan should have a great dependency on the a/c speed too... instead it doesn't.... why?
The aerodynamic profile of a fan blade is slightly different from a propellor. (As they have different design aims)
A propellor is designed to produce most of the forward thrust of an engine whereas a fan blade is designed to be a integral part of the compressor (speeding up the air and compressing it at each succesive stage) . The front fan of a bypass engine is simply the first stage of the compressor albeit large one. This allows a larger portion of the resultant airflow to pass the main core/combustion process
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Originally Posted by soundlover
Why the fan stage is not considered to be the same as a propeller.... the only difference looks to be the number of blades!!!! hence it should behave as a propeller!!!!
PBL
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A propellor is designed to produce most of the forward thrust of an engine whereas a fan blade is designed to be a integral part of the compressor
