Clive,
What am I missing please?
On the nozzle efficiency chart:
If pj goes up, say, ps/pj goes down. Also if pj goes up the primary jet will expand some more and the 2ndary annulus area would decrease.
But 2ndary annulus area decreasing arrow is pointing up. I'd have expected down.
I'm thinking of what's happening at a particular nozzle plane. Perhaps the "2ndary annulus area decreasing" arrow means in the direction of flow?

Some day I'll be smart enough to just barely understand that question, let alone the answer thereto.

flyboyike, I didn't understand the question either after I'd written it. BTW I can barely control my car never mind a plane.

Lyman,

I misunderstood. Don't waste time following up.

peter

US PATENT NO. 3344606, Robert Abernethy

PAGE FIVE, PARAGRAPH THREE, SENTENCE TWO.

"Additional Benefits...."

"....cool air is provided to the afterburner....for cooling purposes..."
( via bleed ducts)

Thanks for your time

:ok:

Peter,

Check your PMs

TD

Clive, for my part I felt unless Lyman could accept that all those associated with the program, including Bob Abernathy, called it a partial ramjet, we would be unable to move forward. How is one able to discuss something if you refuse to acknowledge its existence? Lyman was quite adamant that there is no such thing as a turbo ramjet or partial ramjet.
Just trying to get your attention Lyman
You are correct that the bleed is an either open/closed system. It operates as a function of engine speed (rotor RPM) and CIT (Compressor Inlet Temperature), which has a maximum allowable limit of 427°C. As to a fixed percentage, I’m not so sure. Some detail.

The engine rotor RPM schedule (see attached graph) shows a continual decrease in RPM, starting at a CIT of about 110°C.

At high Mach number and constant inlet conditions, engine speed is essentially constant for all throttle positions down to and including IDLE.
At a fixed throttle position, engine speed will vary when CIT (Mach) changes.

The position of the nozzle is controlled by engine rotor RPM. Throttle movement in the afterburner range will also change nozzle position. At high altitude and maximum afterburner the nozzle will be 80-100% open normally.

As speed (Mach) increases the engine is provided less and less of the overall thrust. It seems intuitive to me that as the inlet, combined with the afterburner, is producing a continually increasing percentage of the thrust, then more and more bleed air, as a percentage, must be being bled off the 4th stage. Hope I make sense.

http://i101.photobucket.com/albums/m...ps33e60098.jpg
http://i101.photobucket.com/albums/m...pse285722b.jpg
http://i101.photobucket.com/albums/m...psc67143d1.jpg

MAXIMUM ALTITUDE CRUISE PROFILE
The maximum altitude cruise profile is 1,000 feet below the maximum afterburner ceiling. Continuous use of maximum afterburner should not be required.

Effect of Mach Decrease
The Mach must not decrease appreciably below the desired cruise Mach. A small decrease in Mach at constant altitude will cause the aircraft to intercept the maximum afterburner ceiling for that speed and become thrust limited. A descent of several thousand feet may he required to re-establish the desired Mach.

Turn Restrictions

NOTE

Turns must be anticipated when flying near maximum altitudes. A descent of approximately 2,000 feet should he completed prior to turn entry.
Use of the maximum altitude cruise and maximum afterburning ceiling profiles is restricted to non turnlng flight. If 35° bank turns are attempted at these altitude schedules, the angle of attack will exceed 8 . Inlet angle of attack biasing will cause compressor inlet pressure to decrease as much as 2 to 3 psi.

Due allowances must be made for the expected altitude loss if maximum power will not be sufficient to maintain level flight.

Effects of Changing Air Temperature
Because of the high true airspeed at cruise, ambient air temperature may change abruptly as different air masses are encountered. Initially, if a constant altitude is maintained, flight into a warmer air mass will cause a decrease in Mach and KEAS, and the true airspeed (TAS) and compressor inlet temperature (CIT] will remain constant. A higher TAS and CIT will result as the desired Mach is reestablished. The opposite would occur as a result of flying into a colder air mass. New cruise altitudes or speeds may be required to compensate for effects of variations in ambient air temperature.

Effect Of Mach Number
For any given gross weight and ambient temperature, the altitudes for maximum range and maximum altitude cruise profiles increase with Mach. This increase is aprox 1,000 feet per .05 Mach. A related characteristic is that if Mach increases slightly above that desired and the throttles are not retarded, excess thrust increases. It is easy to exceed target Mach inadvertently. (peter, this I think is the kernel to your reducing fuel flow question.)

Mach, KEAS, Altitude Relationship
The selection of values for any two of the Mach, KEAS, or altitude variables automatically defines the value of the third, regardless of ambient temperature. For instance, if cruise is scheduled for Mach 3.0 and the desired initial cruise altitude is 72,000 feet, the KEAS must be 396 knots.

Peter,
Ah! I see your problem. My fault - when I edited the original for general consumption I removed some of the labeling. The three lines actually represent three different nozzle designs with varying physical annulus area for a constant primary jet area. The top line has a minimum secondary area 1.66Aj, the middle line 1.83Aj and the lowest line 2.0Aj. The 'aerodynamic' annulus area will of course depend on this and the jet expansion
So when the arrow shows a lower annulus area (pointing up) that is more restrictive of the secondary flow and is consistent with your point - if the secondary flow is squeezed it will require a bigger Ps/Pj to pass a given secondary flow. So when, for a given nozzle geometry, Pj changes the flow characteristic moves along whichever of those three lines is relevant to that nozzle, not across them.
Sorry about that, hope it is clear now

Clive

Lyman,

I definitely do not want to get involved in all that again !

I will just say that nobody has ever claimed that the cooling bleed flow was some sort of ramjet - it cannot be so, if only because there is no burner anywhere in the stream. What others have said is that the designer of that bit of the engine regarded the compressor bleed being fed back into the jetpipe downstream of the turbine and then exposed to afterburning as having some of the features of a ramjet, and he chose to describe the engine as a turbo-ramjet or partial ramjet.

That's fine by me - it was his engine, and as Juliet said "A rose by any other name would smell as sweet"

Brian

I recognised your position - the semantics came from the other side :ok:

Brian
I don't think it can be a fixed percentage. Bleed open/closed is the equivalent of Concorde's secondary air doors (between intake and engine bay) being either open or shut. Once open the bleed flow would be controlled by the flows in the nozzle as explained in my other posts. It would vary with primary jet nozzle area for example, with bigger bleed flows possible at reduced Aj

CliveL

I failed to make it sufficiently clear. I was not in denial about who said "Partial Ramjet" or other language.

I was surprised to see Brian quote the inventor "I called it a Partial Ramjet"...Since I had previously stated "There is no such thing...."

I do not seek to perpetuate unpleasantness, and in my stubborn way, by defending the lack of a connection of "Partial Ramjet" to the actual mechanics of such, I put many people off. I note this, and am regretful.

Thank you for all your help.....

Brian, for the record, I did not claim that a "Turbo Ramjet" does not exist, only that the inventor specifically took the time in his patent to say that his device is NOT a "Turbo Ramjet".

So, I am left wondering WHY the good doctor claimed "partial Ramjet". There is no passive compression, there is no fuel introduced, and there is no "auto-ignition". I'll keep looking.

:ok:

Lyman,
Lyman, why don't you stop this expedition you have been on, no need to continue to keep looking. Dr Bob Abernethy is still alive.

Dr. Bob welcomes questions by email at [email protected].

E-mail him and ask your questions...

TD

I'm not sure what you mean by the term. The inlet increases pressure from an ambient of .4 PSI at 80,000, feet to 18 PSI.

I think there may be a little disconnect in thought as to what constitutes a ramjet.

Taking the engine in isolation does not a partial ramjet make. Asking if a TF30 is a ramjet, the answer is no, though could be if the correct inlet was attached, all other things being considered equal (engine metallurgy etc).

For a ramjet to begin operation, the vehicle has to be accelerated to a speed where the inlet will "start". On the SR-71 the inlet will "start" between 1.6 and 1.8 Mach usually. This is when the normal shock moves from the front of the inlet to a position near the shock trap bleed in the throat.

When above Mach 1.6 the spike will retract approx 1-5/8 inches per 0.1 Mach number. Total motion is approx 26 inches. This increases the captured stream tube area 112%, from 8.7 square feet to 18.5 square feet. The throat closes down to 4.16 square feet, 54% of the area t Mach 1.6.

Have read, starting at page 21 of http://ftp.rta.nato.int/public//PubF...AVT-185-05.pdf

Brian,
Thanks for all the charts. All a bit mystifying.

That's good , thanks. The only value I have ever seen quoted for the cruise condition is 20% . In the patent, in 'Untold tales' and in Peter Law's presentations in AEHS website for example. Never seen a higher value to go with all the other oft-quoted thrust, etc benefits at M3.

What I have just found is the min value when the bleed is opened at M2.
In 'More never told Tales' ... "My solution was to open the valves at a lower Mach number, around 2.0 where there would be no bleed flow and no hiccup." I guess the turbine exit pressure wasn't much lower than the 4th stage exit at that flight speed.

I have more Qs coming.

Brian,
That's good. Thanks.

Also can we see it in terms of reaching its design point ?
ref Col Graham:
"The faster it flew the more efficient it became. For example the range charts show.."

If we plot FF v Mn from the range charts we get a steady increase in FF peaking at M3.0 then a dip to M3.15 and increasing again at M3.2 (for all but one condition).
Isn't this FF trough an indication that the whole aircraft has finally reached its design point. ie it's more efficient at M3.15 than at M3.0 or M3.2?

eg the spike shock doesn't meet the cowl lip until the design speed, the terminal shock is now correctly positioned with minimum intensity, etc.

eg the nacelle drag is a minimum. ref Col graham "Any time the SR-71 was at of above M3.05 the aft bypass was always placed in the CLOSE position."

eg the engine/afterburner/exhaust expansion are all where they should be.

peter, I'll post fuel flow figures for a few different Mach points so you can see the differences. Unfortunately I have a trip on and won't be able to do it until Wednesday at the earliest.

Hi Lyman,
I have woken up in the middle of the night and cannot sleep, which is OK because I'm retired.:ok:

Whether we accept or not that there may exist a particular arrangement of machine called a turboramjet perhaps we can invent a definition for ourselves.

A ramjet consists of a static compressor followed by a combustor followed by a static expander.

There can also be a machine consisting of a static compressor followed by a rotary compressor, or supercharger, coaxial with a gas producer consisting of a rotary compressor, combustor and rotary expander with waste heat recovery. This combination of supercharged air and waste heat recovery is fed, for enhanced combustion efficiency, to a main combustor, also called an afterburner. This is followed by a static expander.

We now have a supercharged ramjet, or, in the rotary compressor/expander vernacular of turbomachinery, a turbocharged ramjet or turboramjet.

Depending on the operating conditions and degree of supercharge, if the static compressor contribution be low and the supercharge be zero, we can call the machine an afterburning turbojet.

May be all rubbish but I did do it in all seriousness because I have a:8 brane.

Peter, no such things as compressors or superchargers, they are called pumps. :E

If you like to send me your home email I'll copy the range charts for you and you can play around deducing what information you wish. 29 charts in all.

Edited to add a few data points.

I've assumed a gross weight of 115,000 pounds and an ISA isothermal atmosphere of -56.5°C. These are for best range.

MachxNautical Miles/1,000 FuelxAltitudexKTAS
x2.4xxxxxxxxxx34xxxxxxxxxxxxxxx63,400
x2.8xxxxxxxxxx37xxxxxxxxxxxxxxx67,000xx1606
x3.0xxxxxxxxxx40xxxxxxxxxxxxxxx70,000xx1721
x3.1xxxxxxxxxx42xxxxxxxxxxxxxxx71,400xx1777
x3.15xxxxxxxxx43.5xxxxxxxxxxxxx72,600xx1806
x3.2xxxxxxxxxx42.5xxxxxxxxxxxxx73,800xx1835

Turbine D, I'm not sure how PMs work so I'm figuring it out and will reply. I thought it meant previous messages but have made one step forward from there so far.
PK

Peter

Hi... I need to repeat. I have not claimed "TurboRamjet" is not real. As you know, Dr. Abernethy describes one in his patent number 3344606 "Recover bleed air system..."

He goes to pains to claim his RBAS, is NOT Turbo Ramjet. i am forwarding a patent that describes a STATIONARY RAMJET, used in power generation.

You might be interested in its function. And its inventor.

Thanks BRIAN, your data is most welcome, I learn alot....

You guys are still going at it? Admirable.

Shhhhhhhhh They're sleeping. Do not disturb!

TurbineD

I received the following via email from Dr. Bob Abernethy this morning....

Lyman,

Thank you for the information you received from Dr. Abernethy.

You say toe-may-toe, I say toe-mah-toe
 

Very good technical thread despite the ad hominid attacks and such, unless they are directed to this old fighter pilot, heh heh. And I think we all learned a lot.

No doubt that the J-58 work helped in the design and performance of the F100 that powered my beloved Viper ten years later. So TNX to Dr Abernathy.

I wonder if we should call that sucker ( the F100) a "turbo-ramjet", as it bypassed a gob of air from the third comperssor stage via an annular duct back to the exhaust/burner ( versus 4th stage in the J-58). Certainly helped the overall thrust when in mil or burner power. Also allowed us a higher Mach than you would expect from a fixed inlet.

One pearl of wisdom from Dr Abernathy's article ( http://www.bobabernethy.com/pdfs/Nev...0of%20P&W3.pdf) has to do with the compressor blade flutter and deformation. So I lost a friend who liked to "run it out" after completing a test hop. Sure enough, one day he got a bit fast (I figure 800 knots indicated or so) and one of the compressor disks went boom and he didn't survive the ejection. Think it was the 4th stage one, not one of the fan ones up front.

All in all, one of the best technical threads I have seen.

hi gums....

I have seen a working sectioned F-100 turning (under electric power) on a stand. It is a marvel.

Abernethy references an annular duct option for the J58, but concludes the cross section allows too much loss of pressure prior to entry into the AB. So he chose the "six big pipes".

I still regret taking such an energetic position v/v "Partial Ramjet"....

A Pratt & Whitney advertisement at Pratt & Whitney, Aviation Pioneers of Groom Lake - Area 51, Nevada quotes bleed bypass ratio as 20 to 40% (above Mach 2 only). The >Mach 2 is understood, as that's when the bleed is scheduled to open.

Brian

If 80% goes through the engine (core), and the rest, 20% the bleeds, what is the percentage of air that bypasses the engine internals by transiting the area between the engine and case (nacelle)? Prior to the Engine intake?

Lyman,

Be careful with the basis of your percentages! 80% of the ENGINE flow goes through the core and 20% of the ENGINE flow through the bleed that feed back to the afterburner. The other bits (shock trap bleed, porous spike bleed etc) should/must be related to INTAKE flow

Hi CliveL. Yes, i thought of that. Here is my problem. The Engine is a "recovery" bleed powerplant. That means that a portion of Engine intake is only temporarily removed, and later re-introduced. So far so good.

How is an accounting of the gaspath accomplished after the reintroduction? The gaspath is still entirely within the confines of the engine Core.

I think Brian has touched on it, but I am interested in the complete inventory of the power profile, by reference to Thrust, not mass. This is the fundamental area I have been trying (badly) to address, and the distillation of what two eminent people claim relative to this complex and amazing engine....

You claim that if related to Bleed recovery, there can be no claim of RamJet.

The inventor claims the opposite, that his Recovery Bleed system is a "Partial Ramjet".

On a technical forum, I woud expect explicit agreement. "Because I say so", has never been a satisfactory reply. I also do not understand the passion relative to this very very minor point.

Thanks for your reply, I remain a BIG FAN.......

Bill

You are mixing things up again! My earlier remarks were specifically directed at the flow that starts at the shock trap bleed and passes over the outside of the engine as cooling air. This is not in any way a ramjet.

The inventor's remarks relate to the internal bleed off the compressor returned to the engine aft of the turbine. Whether or not this can be regarded as a partial ramjet has been thrashed to death already in this thread.

CliveL, noted, and again, thank you for a personal response.

:ok:

Clive,
I've gone through your flow explanation and appreciate it.
I'd like to add, ref
From, for example, "Fast Jets - the history of reheat development at Derby" by Cyril Elliott "... in order to keep the mass flow constant, the area of the nozzle must increase with the sqrt of the jet temperature." (Since gas specific volume has increased significantly).

This book incidentally probably holds the world record for foldout size, 66" for the Adour reheat fuel system schematic.

Hello Lyman,
Don't know if I can help but for an attempt can you be a bit more explicit?
Do you mean, for example, why/how do we get an increase in thrust as a result of the bleed flow becoming available at the afterburner?
Or maybe something else?

Clive,
I've just noticed a Pj/Ps of 0.29 for the SR-71 at M3.2 (Peter Law's presentations on AEHS website). Can that tell us anything by itself on how the SR71 installation compared to Concorde?

Peter
Not by itself Peter, because as you point out the primary jet pipe area will be increased when you light up the afterburner and that will 'squeeze' the secondary flow. Concorde of course operated 'dry' in cruise. That same website (I think) gives the cooling air temperature at the nozzle and the primary jet nozzle pressure and temperature, so if one knew the fully open jet pipe area (can anyone help?) it should be possible to make some sort of comparison.

As it stands Ps/Pj is higher than a typical Concorde value (0.25) so there should be more cooling airflow, but this will be offset by that squeezing effect.

PS I left out the effect of increased Aj in my earlier explanation as I thought it would complicate things :ouch: - that's was why I was careful to specify primary jet area unchanged. I'm guessing that if Abernethy says cooling flow was increased when one lit afterburner than the temperature effect would outweigh the area change ....

New question
 

Where would the intake contribution to the +47% come from? ie not the fundamentals of where intake thrust comes from, but specifically what would have caused the change.

eg would it have been from upsizing the intake to handle the +22% engine flow and hence a greater area of rear-facing surfaces?

Difficult to say for sure without knowing more of the intake characteristics [I'm looking for something that might explain it ]

AS A GUESS, from the fact that it was an installed thrust improvement, it could have been that at Mach 3.0 the intake was actually too big for the engine without bleeds, so that they may have been spilling air through the forward bypass doors (to maintain the normal shock in its correct location). This air was spilled out sideways so no thrust recovery - just momentum drag. This could have been the case if they originally sized the intake assuming the engine would swallow more than it actually would without those bleeds or because the intake was actually sized by some other design case and they had to compromise a little bit. With the increased engine mass flow possible with the returned bleed open, they could keep the forward bypass doors closed.

One more question to add to the list .....

Clive/ Brian,
This is probably too-simplistic a viewpoint to have any merit for such a complex subject, but have to ask.

A lot of interest is always shown in the thrust contribution from the intake at high speeds. For a given installation it goes up with speed, eg for the SR-71 at M2 13%, at M3 54%.

If we now look at Concorde at M2 it's even higher at 63% than the SR-71 at M3.

Both installations were state-of-the-art for their respective design points.

Whilst I understand the actual numbers defining all the flows/thrusts, etc were very different I'm clinging to the idea that a comparison of different installations is perhaps valid using these percentages as they are ratios (non-dimensional).

Question. Is there some fundamental (simple?) reason why Concorde's intake thrust contribution is somewhat higher at a significantly lower speed?

Thanks.

Because it was optimized for M2, and Blackbird for M3.