max altitude and Mach 1

I 've noticed that many of the more modern fighters (F-16, SU-27) only reach
max alt while subsonic, whereas older fighters ( F-4, MiG-21 for instance)
reach max altitude around M 1.6 or more. Is there some general reason for
this?


--



Curiosity killed the cat, and I'm gonna find out why!

In article ,
"Boomer" writes:
I 've noticed that many of the more modern fighters (F-16, SU-27) only reach
max alt while subsonic, whereas older fighters ( F-4, MiG-21 for instance)
reach max altitude around M 1.6 or more. Is there some general reason for
this?

You've asked a Short Question with a Long Answer, I'm afraid.

Evaluating airplane performance, especially from the stuff published
in the Popular Press, is a tricky business. Most of the "Over the
Counter" stuff gives you single data points, and is hopelessly lacking
in context. With airplanes, especially jet fighters, where weight can
vary greatly between Takeoff, Mid-Mission (Over the Target) and
Landings, and there can be many combinations of drag-producing
external stores, it's vital to know what the configuration of the
airplane is at the point that you're measuring. The second problem is
that you don't know the Power Settings being used for these data
points that you're seeing. You can make assumption - but you have to
be careful. The third factor is airframe limitations. Many aircraft
have the thrust available to go faster than they are cleared for in
service. They are limited to lower speeds because of a number of
factors - the ability to withstand gust loads, the strength of the
materials the airframe or powerplant are made of, limitation on the
release of weapons, or loss of stability.

The ceiling of an airplane is the point reached when the Specific
Excess Power in level flight = 0. That means that all the thrust
that's being generated is being used to overcome drag, and there's no
more extra push available to counteract the deceleration produced by
Earth's gravity when you point the flight vector (It's not just the
airplane's nose - the nose-up angle can significant;y differ from the
angle that the airplane is flying.) up. Excess Power depends on two
constantly changing factors - the Amount of thrust that you're able to
generate, which is affected by airspeed, altitude, and Mach Number
(Oh, yeah, Airspeed and Mach Number, while related, aren't the same
thing - Equivalent Airspeed is a measure of the dynamic pressure of
the air as referenced from Sea Level, and is important for determining
how the powerplant and airframe behave. Mach Number (Ratio of current
True Airspeed to the Local Speed of Sound) is all about generating
shockwaves, and how they interact with various parts of the airplane.)

And the amount of Drag that your airframe has, which is controlled by
the airplane's shape, (Drag Coefficient), Weight (To an extent, it's
really not all that important at high speeds in level flight) Mach
Number, and Equivalent Airspeed.

The thrust of a jet engine, whether Afterburning or in dry thrust,
isn't constant. It's greatly influenced by Altitude- Thinner Air =
Less Thrust - and by airspeed in two different ways - Ram Drag, which
is the drag produced by the volume of air the engine needs being
shoved into the inlet, and which reduces thrust linearly with
airspeed, and Ram Compression, which is the fraction of the Dynamic
Pressure (EAS) of the air that's recovered by the inlet system, and
which is basically proportional to the Square of the Airspeed. A
third factor is mechanical limitations of the engine, which affect the
maximum power output possible. These can be materials limitations,
such as the temperature in various parts of the engine, or
performance limits, such as the ability of the fuel pumps to deliver
sufficient fuel to meet the demands placed on the engine.
Ram drag is pretty straightforward - it's just the mass of air
multiplied by the velocity If you can keep the amount of air being
swallowed by the inlet matched to the air required by the engine, it's
not a big problem. This is one function of variable inlets - they
bleed off excess air, and dump it overboard before it can contribute
to Drag. (Another approach is to design an inlet optimized for high
speed (Smaller volume flow) conditions, and have extra inlets that
open up at low speeds to ensure the proper mass flow of air - for
example, the F-105 had fairly small, high speed inlets, but also had
auxiliary air inlets in the main gear wells. When the airplane was
flying slowly, and the main inlets weren't able to supply enough air,
(The landing and takeoff case) the Auxiliary air inlets took up the
slack. You also see it in the translating inlets of, say, the F-111,
which moves forward to expose a slot for more air to enter the inlet,
or the spring-loaded doors around the edges of a Harrier's inlets,
which are sucked open by the increased air demands while hovering an
at low speeds)
Ram Compression is also pretty straightforward, until you start flying
fast enough to generate shockwaves - the Transonic and Supersonic
regions. Shockwaves can really muck up the efficiency of the Pressure
Recovery of an inlet - While an inlet might have a Recovery Coefficient
(The ration of actual pressure recovery to theoretical pressure
recovery) of, say, 0.9 at Subsonic speeds, it may be as low as 0.5 at
Mach 2. This is because getting the air through a strong shockwave
eats up a lot of energy, and so there isn't as much to gain after it's
been through the shock. One solution to that is to pass the air
through a series of weaker shockwaves, each of which takes a smaller
chunk of the energy from the incoming air. This is done by balancing
these shocks against each other, and allowing excess air to spill
out. A good multi-shock inlet can have a Recovery Coefficient of 0.85
- 0.9 at Mach 2.
That sounds simple enough, but these shockwaves are very dynamic
things, changing their angles and strengths as the Mach Number
changes. That's why many jets intended for high-speed flight have
variable inlets - they keep the ram drag to a minimum, while ensuring
the optimum pressure recovery. (When they work right - it was a big
problem in the early days)
So, the upshot of all this is that the Thrust of a jet engine
decreases a bit as it starts moving through the air, (Ram Drag), then
starts increasing (Pressure Recovery) until some limit, such as
Incoming Air Temperature (As you compress the air in the inlet, it
gets hotter - as you compress it in the engine, it gets hotter still -
at some point, you either will melt something in the engine, or not be
able to add enough energy by burning fuel to produce more thrust.)
or the ability of the fuel system to provide enough fuel to heat the
volume of incoming air to the level required to produce more thrust.

The non-afterburning parts of the engine (Core/Gas Generator) are
generally limited by material temperatures. This makes all engine
design a compromise - A high pressure ration in the compressor section
of the engine is great for fuel economy in dry thrust, but because of
the added temperature increase in the compressor, it reached the
material limits sooner. A low Pressure Ratio allows for more Ram
Compression, and is more efficient (non-afterburning) at high speeds,
but drinks a river of fuel. Afterburning thrust isn't much different,
however. The Fuel Control on the engine is heating the air in the
afterburner section to the same temperature either way - no serious
materials limitations, and the thrust produced in AB, and the amount of
increase with airspeed, is about the same. (In this case, fuel
consumption for the High Compression Ratio engine can be a bit higher,
because the hot gas coming from the Gas Generator is a bit cooler)

The final upshot of all this is that the thrust increases with speed,
and at Mach 2 can be more than twice as much as it is when not moving,
but is affected by a number of factors.

Now we come to Drag. At subsonic speeds, it's pretty simple - you've
got some measure of how much the airplane resists moving through the
air, which is based on the size of the airplane and its shape and
finish, and the Dynamic Pressure of the air at the speed the airplane
is going, (EAS). When you start getting near the Speed of Sound,
though, it gets complicated - the speed of the air moving past a body
isn't constant, and it accelerates as it goes past Curvy Bits, like the
side of the fuselage, the top and bottom surfaces of the wings, and
any stuff stuck to the basic shape, like Canopies, air scoops, bombs,
missiles, and the like. (It also slows down at other points.
Transonic and Supersonic Aerodynamics can be counter-intuitive and
Ugly). Drag starts showing up in ways that weren't there before -
there's Wave Drag from the wings, Form Drag and Base Drag on the
fuselage, and it gets Really Ugly. It starts to settle out at about
Mach 1.4 or so, as the airflow over the entire airplane becomes
supersonic, but it's still complicated.
Basically, the Drag Coefficient (The measure of how slippery the shape
is) of a wing will start to increase around Mach 0.7 - 0.8 or so,
depending on the airfoil thickness (Thin wings are good for going
fast, but don't work well going slow, and are hard to build strong -
tradeoffs, again). The Drag Coefficient of a wing increases sharply
(About 10X, for a unswept wing) at Mach 1, and then drops off.
This can be changed by sweeping back the wing. That "fools the air"
into thinking that the wing is moving slower than it is, and can
somewhat raise the speed at which the Drag Rise starts, raise the Mach
Number where the Drag Peak occurs, and limit the amount of the drag
increase. But it can't make it go away.
The fuselage, usually being a longer, narrower shape, isn't as
affected. But at around Mach 0.85 or so, the drag starts to increase
there, as well, and peaks at a bit over Mach 1 in a normal case. It
decreases some after that, usually, but it's generally around 1.5 times
the subsonic drag. The amount of this increase is influenced very
much by the overall shape of the airplane - basically, the most
efficient transonic/supersonic shape is that of a rifle bullet - a
pointed nose, and a length about 12 times longer than the diameter,
with a smooth sweeping curve along the length. That doesn't happen in
real life, though - The wings, engines, canopy, tail, and all those
other bits sticking out disturb both the ideal shape, and the
smoothness of the curve. This can be alleviated somewhat by allowing
for the change in cross section along the length of the airplane -
tucking some bits in, and bulging others to make the aggregate shape
more closely match the ideal distribution. That's the Area Rule.

Anyway - to the Short Form - The Drag Coefficient of an airplane stays
fairly low until about Mach 0.8. It then increases in an amount
dependant on the shape of the airplane to a point somewhere over Mach
1, and can increase by a huge amount, the peak Mach Number, and the
amount depending on the particular shape of the airplane at that
moment. (As you can see from the previous paragraph, external stuff
like bombs, drop tanks, or missiles can really muck things up) the
Drag Coefficient than usually drops a bit, but it's never as low as it
is in the subsonic case.
That's just the Drag Coefficient. The actual drag is the product of
the Drag Coefficient, the size of the airplane, and the Dynamic
Pressure of the air that it's moving through, which increases with the
Square of the velocity. This means that Drag is always increasing,
and sometimes, die to the change in drag coefficient, can be
increasing sharply.

So, we end up with a situation where Thrust is increasing due to
velocity, and Drag is increasing in velocity at a slightly greater
rate. With the lumpy shape of the Drag Curve, that usually means that
there's a peak in Excess Power (and thus ceiling) at about Mach 0.9.
Excess Power drops as all the transonic Drag Rise stuff occurs, and
after that, the Thrust and Drag increase at almost the same rate, with
a bit of a peak for Thrust somewhere around Mach 1.7.
If you can actually get a copy of the 1G flight envelope of a
supersonic airplane. (The Standard Aircraft Characteristics Charts are
a good place to look) you'll see that there's usually a peak in the
ceiling at about Mach 0.9, and another at about Mach 1.7.
This of course, is only a rule of thumb - some airplanes are a bit
different.
The F-16, for example, has an airframe optimized for peak
Excess Power in the Mach 0.9 area, where most air-to-air combat
currently occurs. It has (well, had, -16s have gotten pretty heavy,
of late) a huge amount of excess thrust at Mach 0.9, and a fixed, more
or less single-shock inlet. Because it has so much excess thrust, it
may still reach a maximum speed in the Mach 2.0 range despite the
lower pressure recovery of the inlet system. But that biases the
Excess Thrust curve in the subsonic direction.

Another example is the F-104A. The USAF flew two versions of the
airplane - the original model had a J79-3 turbojet, which produced
about 9600 lbf of thrust without Afterburner, and roughly 15,000 lbf
of thrust with the afterburner operating. (All values Static Sea
Level) and 2-shock inlets. T
he performance pretty much matching the Rule of Thumb stated above.
With a Ceiling of about 50,000' ad Mach 0.9, 55,000' at Mach 1.7. The
airplane was limited structurally to 750 KEAS (Knots Equivalent
Airspeed), and an engine inlet temperature of 250 Deg F (Roughly Mach 2).
It generally could reach the 750 KEAS limit between 20,000' and
35,000', and the 250 Deg F limit was reached between 35,000' and a
shade over 50,000'.

In the mid '60s, wanting a higher performance Interceptor for Southern
Florida, the USAF re-engined some of the F-104As with the J79-19
engine used on late model Phantoms. This had a non-afterburning
Static Thrust of 11,900#, and an Afterburning Static Thrust of
17,900#. With that much power, the 750 KEAS airspeed limit was
reached at all altitudes, from Sea Level on up, and the 250 Degree F
limit was reached from 20,000' to the maximum ceiling of around
66,000'. The ceiling continuously increased from 51,00' at Mach o.9
to 66,000' at Mach 2.0. It could very easily have flown higher and
faster, if the airframe limits were ignored.

Sorry for the long answer. That sometimes happens with short
questions.

--
Pete Stickney
A strong conviction that something must be done is the parent of many
bad measures. -- Daniel Webster

Sorry for the long answer. That sometimes happens with short
questions.



Maybe you could answer a question I've had for a long time. If you
look at the SR-71's inlets from the side they seem to be pointing
somewhat down. I took this to mean that since it seems you'd have to
have the inlet lip on a circular inlet perpendicular to the airflow to
maximize it's efficiency, that at cruise speed and altitude the
Blackbird would be flying at an angle of attack such that the inlet
lip would be at 0 degrees AOA. At that angle the exhaust would exit
in a somewhat downward direction. So my question is is that setup to
maximize the altitude potential (because thrust would be directly
aiding lift)? Do ALL aircraft fly at a certain angle of attack at
their maximum altitude? Is the only reason you see these things on
the Blackbird because it's designed to spend most of it's time in
those conditions? Would a Blackbird's max altitude also be at Mach
0.9?

Thanks.

Pete S., I dont mind long answers at all, I will in all likelyhood read your
message a number of times and glean every spec of knowledge I can from it.

--



Curiosity killed the cat, and I'm gonna find out why!
"Scott Ferrin" wrote in message
...


Sorry for the long answer. That sometimes happens with short
questions.



Maybe you could answer a question I've had for a long time. If you
look at the SR-71's inlets from the side they seem to be pointing
somewhat down. I took this to mean that since it seems you'd have to
have the inlet lip on a circular inlet perpendicular to the airflow to
maximize it's efficiency, that at cruise speed and altitude the
Blackbird would be flying at an angle of attack such that the inlet
lip would be at 0 degrees AOA. At that angle the exhaust would exit
in a somewhat downward direction. So my question is is that setup to
maximize the altitude potential (because thrust would be directly
aiding lift)? Do ALL aircraft fly at a certain angle of attack at
their maximum altitude? Is the only reason you see these things on
the Blackbird because it's designed to spend most of it's time in
those conditions? Would a Blackbird's max altitude also be at Mach
0.9?

Thanks.

In article ,
Scott Ferrin writes:


Sorry for the long answer. That sometimes happens with short
questions.



Maybe you could answer a question I've had for a long time. If you
look at the SR-71's inlets from the side they seem to be pointing
somewhat down. I took this to mean that since it seems you'd have to
have the inlet lip on a circular inlet perpendicular to the airflow to
maximize it's efficiency, that at cruise speed and altitude the
Blackbird would be flying at an angle of attack such that the inlet
lip would be at 0 degrees AOA. At that angle the exhaust would exit
in a somewhat downward direction. So my question is is that setup to
maximize the altitude potential (because thrust would be directly
aiding lift)? Do ALL aircraft fly at a certain angle of attack at
their maximum altitude? Is the only reason you see these things on
the Blackbird because it's designed to spend most of it's time in
those conditions? Would a Blackbird's max altitude also be at Mach
0.9?

Well, I'll try - Yes, an A-12/YF-12/SR-71's inlets do face down a bit,
and the reason is to present an inlet face perdenicular to the
airflow, as much as possible. The Blackbirds were intended to cruise
right at the edge of what was possible for an airplane that could also
take off & land, back in the late 1950s. They needed to squeeze every
mit of efficiency out of the airframe & powerplant (Which can be hard
to tell apart, on an SR), and the airplane was intended to get itself
to one point in its flight envelope and stay there. (Mach 3.2/80,000'
or so, around 375 KEAS) At that EAS, an for teh weights that would be
expected, the Angle of Attack range would be predictable, and so it
was dialled in to the inlet design. This maximizes the inlet
efficiancy, and helps alleviate the possibility of the inlet getting
dicombobulated with the complex series of shock waves that it uses to
allow for the maximum pressure recovery. Consider how much of a
problem inlet "unstarts", where the shocks got all tangled up & the
inlet system stopped properly supplying air to the engine, were in the
early stages of the program. Then think about how much worse it would
have been with the inlets getting an uneven flow. Very Ugly Indeed.

While thrust vectoring with AoA does occur, (A good example would be
the F-104. I was told by a CanForce CF-104 pilot that the best way to
ensure a hard landing was to pull back on the throttle during the
flare - the AoA was high enough that a fair chunk of hte airplane's
weight was riding on teh vertical component of the thrust), I don't
think that that was a factor. The angle's too small for there to be
much of a vertical component to the thrust. It might have an effect
on cruise trim, though.

--
Pete Stickney
A strong conviction that something must be done is the parent of many
bad measures. -- Daniel Webster

Peter Stickney wrote:

snip

While thrust vectoring with AoA does occur, (A good example would be
the F-104. I was told by a CanForce CF-104 pilot that the best way to
ensure a hard landing was to pull back on the throttle during the
flare - the AoA was high enough that a fair chunk of hte airplane's
weight was riding on teh vertical component of the thrust), I don't
think that that was a factor.

Pete, that would also lose the BLC, which was almost certainly a larger
effect. The Dash-1 says

"The pilot should remember at all times when using LAND flaps that the
additional lift afforded by BLC is dependent on engine airflow. This lift,
therefore, varies with airspeed, altitude, and engine rpm with the greatest
variation occurring at low airspeed, low altitude, and engine speeds above
80%. This means that the proper use of the throttle is mandatory,
especially as touchdown is approached, to accomplish a smooth reduction in
engine rpm so that a smooth reduction in the effects of BLC will result."

Guy

The throttle reduction restriction in the F-104 was to ensure that the BLC
airflow was maintained over the rear flaps until the aircraft touched down.
The minimum RPM as I recall was 83%. In any case, the F-104 did not fly at a
high angle of attack with full flaps down...it's pitch attitude was much
"flatter' than other fighters of that period that did fly at a high AOA (the
F-4, for example).
"Peter Stickney" wrote in message
...
In article ,
Scott Ferrin writes:


Sorry for the long answer. That sometimes happens with short
questions.



Maybe you could answer a question I've had for a long time. If you
look at the SR-71's inlets from the side they seem to be pointing
somewhat down. I took this to mean that since it seems you'd have to
have the inlet lip on a circular inlet perpendicular to the airflow to
maximize it's efficiency, that at cruise speed and altitude the
Blackbird would be flying at an angle of attack such that the inlet
lip would be at 0 degrees AOA. At that angle the exhaust would exit
in a somewhat downward direction. So my question is is that setup to
maximize the altitude potential (because thrust would be directly
aiding lift)? Do ALL aircraft fly at a certain angle of attack at
their maximum altitude? Is the only reason you see these things on
the Blackbird because it's designed to spend most of it's time in
those conditions? Would a Blackbird's max altitude also be at Mach
0.9?

Well, I'll try - Yes, an A-12/YF-12/SR-71's inlets do face down a bit,
and the reason is to present an inlet face perdenicular to the
airflow, as much as possible. The Blackbirds were intended to cruise
right at the edge of what was possible for an airplane that could also
take off & land, back in the late 1950s. They needed to squeeze every
mit of efficiency out of the airframe & powerplant (Which can be hard
to tell apart, on an SR), and the airplane was intended to get itself
to one point in its flight envelope and stay there. (Mach 3.2/80,000'
or so, around 375 KEAS) At that EAS, an for teh weights that would be
expected, the Angle of Attack range would be predictable, and so it
was dialled in to the inlet design. This maximizes the inlet
efficiancy, and helps alleviate the possibility of the inlet getting
dicombobulated with the complex series of shock waves that it uses to
allow for the maximum pressure recovery. Consider how much of a
problem inlet "unstarts", where the shocks got all tangled up & the
inlet system stopped properly supplying air to the engine, were in the
early stages of the program. Then think about how much worse it would
have been with the inlets getting an uneven flow. Very Ugly Indeed.

While thrust vectoring with AoA does occur, (A good example would be
the F-104. I was told by a CanForce CF-104 pilot that the best way to
ensure a hard landing was to pull back on the throttle during the
flare - the AoA was high enough that a fair chunk of hte airplane's
weight was riding on teh vertical component of the thrust), I don't
think that that was a factor. The angle's too small for there to be
much of a vertical component to the thrust. It might have an effect
on cruise trim, though.

--
Pete Stickney
A strong conviction that something must be done is the parent of many
bad measures. -- Daniel Webster

You've asked a Short Question with a Long Answer, I'm afraid.

Peter, why do I suspect that when asked the time, you tell the inquirer how
to build a watch?

Evaluating airplane performance, especially from the stuff published
in the Popular Press, is a tricky business.

Big snip.

In the mid '60s, wanting a higher performance Interceptor for Southern
Florida, the USAF re-engined some of the F-104As with the J79-19
engine used on late model Phantoms. This had a non-afterburning
Static Thrust of 11,900#, and an Afterburning Static Thrust of
17,900#. With that much power, the 750 KEAS airspeed limit was
reached at all altitudes, from Sea Level on up, and the 250 Degree F
limit was reached from 20,000' to the maximum ceiling of around
66,000'. The ceiling continuously increased from 51,00' at Mach o.9
to 66,000' at Mach 2.0. It could very easily have flown higher and
faster, if the airframe limits were ignored.

Nice try, but untrue. The 750 airframe limit was not a factor above about
40,000 feet ... it was not reached at "all altitudes." (BTW, airframe limit
IS a factor ... was? ... for the SR-71 at intermediate altitudes.) Inlet
temperature could be an issue at the extreme top end ... Skyburner F-4 had
and Greenameyer's F-104 was to have inlet water injection ... but we're
talking 2.5 plus here.

As to "very easily flown higher and faster" the J-79 would experience burner
blow out between 65-70,000 feet and the engines would have to be shut down
approaching 75,000 because their minimum fuel flow settings would be too
high and cause overtemp. (Greenameyer intended to modify the fuel control
and use specially formulated fuel to allow the engine to run longer until
shutdown required in his zoom climb.)

To simplify your response, most older designs had high mach as a primary
design goal and thrust/drag created large PsubS "bubbles" past the transonic
drag rise region (F-104 a prime example, original F-14B ... glove vanes and
inlet scheduling intact ... another). That excess power in the 1.4-1.6
region (usually, SR-71 was much higher) allowed higher service ceilings
while supersonic.

Current design emphasis is on subsonic performance with high Q (indicated
airspeed) but not usually high mach as a bonus of their high thrust/weight
ratios. No large PsubS gains once above transonic drag rise. Ergo no
improvement in service ceiling supersonic.

R / John

In article ,
"John Carrier" writes:
You've asked a Short Question with a Long Answer, I'm afraid.

Peter, why do I suspect that when asked the time, you tell the inquirer how
to build a watch?

Evaluating airplane performance, especially from the stuff published
in the Popular Press, is a tricky business.

Big snip.

In the mid '60s, wanting a higher performance Interceptor for Southern
Florida, the USAF re-engined some of the F-104As with the J79-19
engine used on late model Phantoms. This had a non-afterburning
Static Thrust of 11,900#, and an Afterburning Static Thrust of
17,900#. With that much power, the 750 KEAS airspeed limit was
reached at all altitudes, from Sea Level on up, and the 250 Degree F
limit was reached from 20,000' to the maximum ceiling of around
66,000'. The ceiling continuously increased from 51,00' at Mach o.9
to 66,000' at Mach 2.0. It could very easily have flown higher and
faster, if the airframe limits were ignored.

Nice try, but untrue. The 750 airframe limit was not a factor above about
40,000 feet ... it was not reached at "all altitudes." (BTW, airframe limit
IS a factor ... was? ... for the SR-71 at intermediate altitudes.) Inlet
temperature could be an issue at the extreme top end ... Skyburner F-4 had
and Greenameyer's F-104 was to have inlet water injection ... but we're
talking 2.5 plus here.

John, where did I say that it was? Ah, never mind, I see where I
didn't state it clearly. Sorry about that. I thought I'd mentioned
that the F-104 flight limits were 750 KEAS or 250 Def F at teh
compressor face, whichever came first. In the case of the -19
powered F-104A, it would run out to the 750 KEAS limit from Sea Level
on up, and the 250 Deg F limit would be reached at anout 20,000', at
about Mach 1.70. Obviously, you'd hold to whichever limit came
first. From 20 Kft on up, the limit you'd run into first was the 250
Deg F limit. At 35,000', the 250 Drg F limit is about 650 KEAS. At
40,000', it's about 550 KEAS. at 50 Kft, it's about 450 KEAS. (BTW,
the SR-71's Q (EAS) limit is fairly low, something like about 450
KEAS.)

Sure- above the Tropopause, the temperature remains constant. and the
250 Deg F limit is reached at about Mach 2.0. If you've got a way to
cool the inlet air as it's being compressed, such as the
Pre-Compressor Cooling on the Skyburner F4H (Pre 1962, after all), or
the similar Water Injection system that Darryl Greenameyer was going
to use, then you can run out to a higher speed safely.


As to "very easily flown higher and faster" the J-79 would experience burner
blow out between 65-70,000 feet and the engines would have to be shut down
approaching 75,000 because their minimum fuel flow settings would be too
high and cause overtemp. (Greenameyer intended to modify the fuel control
and use specially formulated fuel to allow the engine to run longer until
shutdown required in his zoom climb.)

That would be higher and faster at the same time - One very
interesting bit from the F-104A (-19) engine's SAC Chart, Jun 1970,
(If you need to see it, I'll be glad to E-mail you a copy)
Is that the ceiling is increasing as it approaches Mach 2,0/66,000'.
That's about 320 KEAS. As far as the engine is concerned, it's
being delivered 320 Kt/Sea Level conditions from teh inlet. They sure seem
to run O.K. in that range. Of course, if you're slower, it'll be a
_lot_ different. But that's the point - With the -19 engines F-104A,
it had the power to go a lot faster than its flight limits would
allow. So it had the potential to, if you were ignoring the limits,
deliver some astounding performance.

There didn't seem to be that much problem with a J79 above 60 Kft -
the B-58 on a high altitude bomb run at Mach 2.0 would be over the
target at 64,000'.

The Rutkowski trajectory for the F-4 and F-104 zoom climbs is fairly
similar - Take off, Accelerate to Mach 0.9, climb at Mach 0.9 t0 a bit
above 36,000' (The Tropopause, where the margin of Thrust over Drag will be
greates, enter a slight descent to get through the transonic drag rise
quickly, accelerate out at 36,000' to however fast you can go, then a
2G pull to straight up and maintain 90 degrees nose high. The
airplane will be decelerateing from that point on, and at 60-70,000'
will be flying at a rather low EAS - somewhere arount 100 KT EAS
wouldn't be too out of line.


To simplify your response, most older designs had high mach as a primary
design goal and thrust/drag created large PsubS "bubbles" past the transonic
drag rise region (F-104 a prime example, original F-14B ... glove vanes and
inlet scheduling intact ... another). That excess power in the 1.4-1.6
region (usually, SR-71 was much higher) allowed higher service ceilings
while supersonic.

That's true - Speed was everything in the '50s, and they found that
with proper inlet design, they could maintain a sufficient thrust
margin to get the airplane up to some big mach number. Let's not
forget too, that Ps is Specefic Excess Acceleration * True Airspeed,
either - the faster you're going, the less excess thrust you'll need
for a given Rate of Climb.

Current design emphasis is on subsonic performance with high Q (indicated
airspeed) but not usually high mach as a bonus of their high thrust/weight
ratios. No large PsubS gains once above transonic drag rise. Ergo no
improvement in service ceiling supersonic.

Right.

Thanks for pointing out my poor wording, above - It seems that we were
talking about the same thing, I just expressed it poorly.

--
Pete Stickney
A strong conviction that something must be done is the parent of many
bad measures. -- Daniel Webster

Peter Stickney wrote:

In article ,
"John Carrier" writes:

snip

As to "very easily flown higher and faster" the J-79 would experience burner
blow out between 65-70,000 feet and the engines would have to be shut down
approaching 75,000 because their minimum fuel flow settings would be too
high and cause overtemp. (Greenameyer intended to modify the fuel control
and use specially formulated fuel to allow the engine to run longer until
shutdown required in his zoom climb.)

That would be higher and faster at the same time - One very
interesting bit from the F-104A (-19) engine's SAC Chart, Jun 1970,
(If you need to see it, I'll be glad to E-mail you a copy)
Is that the ceiling is increasing as it approaches Mach 2,0/66,000'.
That's about 320 KEAS. As far as the engine is concerned, it's
being delivered 320 Kt/Sea Level conditions from teh inlet. They sure seem
to run O.K. in that range. Of course, if you're slower, it'll be a
_lot_ different. But that's the point - With the -19 engines F-104A,
it had the power to go a lot faster than its flight limits would
allow. So it had the potential to, if you were ignoring the limits,
deliver some astounding performance.

There didn't seem to be that much problem with a J79 above 60 Kft -
the B-58 on a high altitude bomb run at Mach 2.0 would be over the
target at 64,000'.

Walt Bjorneby must be busy, or I'm sure by now he'd have mentioned his cruising in
his F-104A w/-19 from Tyndall to Homestead at M2.0 and FL730 (he'd filed IFR at
1,120 KTAS and that altitude). I believe he said he was using about 3/4 AB and
burning 6,000 pph.

Guy