Forum: F-22A Raptor

Last edited by dwightlooi on Mar 30, 2008 - 07:07 PM; edited 2 times in total

There has been some discussion about the feasibility of marrying the ESSM's 10" motor with the AIM-120's 7" front end to create a long range AAM for the F-22. This is of particular interest because six will fit neatly in the F-22A's belly bays with more room to spare than with the AMRAAM.

Let's analyze this proposition:-

Fit -- ESSM based solution fits with space to spare.

AMRAAM = 3.65 m (longitudinally)
AMRAAM = 12.5" (vertical depth)
3 x AMRAAMs occupies 2.75+7+2.75+7+2.75+7+2.75 = 32" (laterally)

ESSM = 3.66 m (longitudinally)
ESSM = 10" (vertical depth)
3 x ESSMs occupies 10+10+10 = 30" (laterally)

Weight -- approximately 550 lbs

ESSM = 620 lbs
AIM-7P front end = ~220 lbs
Mk134 10" Motor = ~ 400 lbs (including tail control section and strakes)

AMRAAM = 335 lbs
AMRAAM WPU-6/B motor = 157 lbs (published weight;Raytheon)
AMRAAM finnage & tail control assembly = ~30 lbs
AMRAAM Front end w/warhead = ~148 lbs (335 - 157 lbs - 30 lbs)

AMRAAM-ESSM Hybrid = 400+148 = ~550 lbs
Propellant Weight = ~320 lbs (est. 80% of 10" motor section weight)
Propellant Fraction = ~58%

Performance -- Pretty darn interesting!

Assumptions:
IpSec of motor = 250 seconds
Drag-Coefficient = ~0.7 (w/frontal reference area; big fin hobby rockets ~0.75)
Frontal Reference area = ~0.06 sq-m (0.051 for 10" body + 0.009 for fins)
Air Density at 12,000 m = 0.232 kg/m^3
Speed of Sound at 12,000 m = 300 m/s

Gross Delta V = 10 x 250 x LN(550/(550-320)) = 2180 m/s = ~Mach 7.3

In an all boost configuration the missile may reach about Mach 7.3 (gross) above launch velocity. Actual burnout velocity will vary depending on trajectory -- hence air drag or fractional gravitational acceleration/deceleration.

Drag at Mach 3 (Newtons) = 0.5 x P x V^2 x Cd x A = 0.5 x 0.232 x 900^2 x 0.7 x 0.06 = 3,946 N = ~887 lbs

With 3/5 boost + 2/5 sustain propellant graining:-

Gross Delta V (@ end of boost phase) = 10 x 250 x LN(550/550-192) = 1,073 m/s = ~Mach 3.6

Approximate post boost velocity (w/Mach 2.0 release) = ~Mach 5.1+ (Delta V corrected at 88% factor for drag loss high altitudes)

Approximate sustainer burn time (@ 887 lbs thrust) = 128 x 250 / 887 = ~36 seconds.

OK, so... as an all boost missile this weapon is estimated to be kinematically capable of going up to Mach 8.4+ on a high altitude ballistic trajectory with a Mach 2 release.

Alternatively, it can also be grained to be capable of reaching Mach 5.1+ and keeping its sustainer lit to overcome ~Mach 3 drag levels for about 36 seconds. Mach 3.5+ is the velocity the weapon will reach with a very low speed launch or a very low altitude transonic release. With a high speed launch such a boost sustain weapon with go to about Mach 5.1 and slowly decelerate while always staying above Mach 3 for the duration of the sustainer burn. Total motor time (boost + sustain) will be about ~45 seconds.

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Now use NCADE for the front end and go after airplanes with it.

Pretty darn interesting, Dwight. Very darn, actually. A fair proposition.

Forum Veteran Joined: Apr 16, 2005 - 03:16 PM Posts: 800

dwightlooi,

I'm not sure about your calculation:

Quote:

Gross Delta V = 10 x 250 x LN(620/(620-320)) = 1815 m/s = ~Mach 6

Why don't you take 550 instead of 620?

That would give about Mach 7.2

Also:

Quote:

Gross Delta V (@ end of boost phase) = 10 x 250 x LN(620/620-192) = 926 m/s = ~Mach 3.1

With 550 instead of 620, that would give M3.57.

Also the missile could be redesigned so that the the maximum amount of fuel can be carried. The AMRAAM front end would be repackaged for a 10'' container. Let's say that would save enough space for 100 more lbs of propellent ( total 420lbs ).

Missile weight would be around 650lbs and propellent weight 420lbs.

In an all boost configuration, that would give a delta V = 10 x 250 x LN(650/(650-420)) = 2597m/s = M8.65

In a boost-sustain combination, a boost with 200lbs of propellent would give a Gross Delta V (@ end of boost phase) = 10 x 250 x LN(650/(650-200)) = 919 m/s = M3.06

and a sustain time of:
220 x 250 / 887 = ~62 seconds!

Also the missile radar would be bigger. And the missile could be lengthened a bit to carry even more fuel, because the missiles would not need staggering.

Let's say that the missile is lengthened, and another 50 more lbs of fuel can be carried for a total of 470lbs and that missile weight is increased by 65lbs.

In an all boost configuration, that would give a delta V = 10 x 250 x LN(665/(665-470)) = 3066 m/s = M10.22!!

In a boost-sustain combination, a boost with 210lbs of propellent would give a Gross Delta V (@ end of boost phase) = 10 x 250 x LN(665/(665-210)) = 948 m/s = M3.16

and a sustain time of:
260 x 250 / 887 = ~73 seconds!

Viperalltheway wrote:

dwightlooi,

I'm not sure about your calculation:

Quote:

Gross Delta V = 10 x 250 x LN(620/(620-320)) = 1815 m/s = ~Mach 6

Why don't you take 550 instead of 620?

That would give about Mach 7.2

Also:

Quote:

Gross Delta V (@ end of boost phase) = 10 x 250 x LN(620/620-192) = 926 m/s = ~Mach 3.1

With 550 instead of 620, that would give M3.57.

You are absolutely right. I am a retard and plugged in the wrong number in the Delta V equation. After going through the trouble to estimate the masses of the AMRAAM/ESSM hybrid, I plugged in ESSM numbers.

Anyway, it has been corrected.

Viperalltheway wrote:

Let's say that the missile is lengthened, and another 50 more lbs of fuel can be carried for a total of 470lbs and that missile weight is increased by 65lbs.

In an all boost configuration, that would give a delta V = 10 x 250 x LN(665/(665-470)) = 3066 m/s = M10.22!!

In a boost-sustain combination, a boost with 210lbs of propellent would give a Gross Delta V (@ end of boost phase) = 10 x 250 x LN(665/(665-210)) = 948 m/s = M3.16

and a sustain time of:
260 x 250 / 887 = ~73 seconds!

Yes... if you do all that, you'll end up with a propellant fraction that is a little HIGHER than that of the THAAD missile. If you do that about 70% of the weapon will be the motor (by length); THAAD is a little less than that (about 63% by length, although the payload is somewhat light). THAAD reaches an altitude of up to ~150km in engagements and burns out at about Mach 8.3+ -- that is from a ground launch with zero initial velocity.

BTW, the THAAD booster has a 17 second all-boost burn profile. Burnout velocity is ~2700m/s although this can be less if the shot is cross range with more time spent at lower altitudes.

src: http://www.princeton.edu/~globsec/publi ... 51-202.pdf

Just a little illustration...

dwightlooi wrote:

BTW, the THAAD booster has a 17 second all-boost burn profile. Burnout velocity is ~2700m/s although this can be less if the shot is cross range with more time spent at lower altitudes.

src: http://www.princeton.edu/~globsec/publi ... 51-202.pdf

I'd be surprised if it's only 17 as it appears to be a longer burning, lower thrust motor based on launch videos. Compare it to SM-3 which weighs more, takes off faster, and whos first two stages alone burn for over 30 seconds.

edit: just finished reading that document. Not sure I buy a lot of what they're saying. First off it's unlikely that the real motor performance is public knowledge as that would give away a lot about the missile's potential performance. I don't doubt that numbers are out that but I do doubt their accuracy. Secondly, they calculate the fuel capacity of the KV by going off the placeholder graphics of a powerpoint presentation?????? YGBSM. While I think the article was an interesting read it's nothing more than another academic WAG.

dwightlooi wrote:

Just a little illustration...

and the real deal. . .

I am thinking more like this...

LEAP would not fit on that.

sferrin wrote:

LEAP would not fit on that.

Yes, it will... the nose fairing just won't be 10" but 13.5 ".

dwightlooi wrote:

sferrin wrote:

LEAP would not fit on that.

Yes, it will... the nose fairing just won't be 10" but 13.5 ".

That is not what you've shown. It will not fit on what you've shown.

Forum Veteran Joined: Apr 16, 2005 - 03:16 PM Posts: 800

dwightlooi,

Wouldn't there be a way to determine at launch the amount of fuel to burn for the boost phase and the amount of fuel to burn for the sustained phase?

For instance if an F-22 launches at 60000fts, since at this altitude the air density is half of what it is at 40000fts ( 12000 m), drag would be cut in half, so the sustained thrust could be maintained twice as long ( 148 sec ).

Also isn't it possible to also determine at launch time the amount of fuel to accelerate to a certain speed, say Mach 4.5, depending on the speed of the aircraft?

And finaly, wouldn't it possible to determine how to use the propellent as best as possible.. If say the target is 50 km away, would it be possible to burn the propellent faster to reduce the duration of the flight.

Viperalltheway wrote:

dwightlooi,

Wouldn't there be a way to determine at launch the amount of fuel to burn for the boost phase and the amount of fuel to burn for the sustained phase?

For instance if an F-22 launches at 60000fts, since at this altitude the air density is half of what it is at 40000fts ( 12000 m), drag would be cut in half, so the sustained thrust could be maintained twice as long ( 148 sec ).

Also isn't it possible to also determine at launch time the amount of fuel to accelerate to a certain speed, say Mach 4.5, depending on the speed of the aircraft?

And finaly, wouldn't it possible to determine how to use the propellent as best as possible.. If say the target is 50 km away, would it be possible to burn the propellant faster to reduce the duration of the flight.

No, it will not be possible. Solid rockets are "throttled" by molding the inner surface geometry of the solid propellant fill. The burn rate -- hence thrust and fuel consumption -- is directly proportional to the internal surface area of the propellant fill. A star shaped hollow burns faster than a hollow cylinder which burns faster than a solid cylinder lit on one end. Most boost sustain type motors have a portion of the propellant shaped with a star shaped center and the rest having a combination of cylindrical and segmented end burn grains.

In other words, you can have just about any reasonable thrust profile you want as long as you set that during the the time of manufacture. Once the motor is made, it is made. It'll burn the way it was designed and made to burn. You can't throttle it like you can a liquid fueled rocket or to a lesser degree solid fuel ramjets.

dwightlooi wrote:

Viperalltheway wrote:

dwightlooi,

Wouldn't there be a way to determine at launch the amount of fuel to burn for the boost phase and the amount of fuel to burn for the sustained phase?

For instance if an F-22 launches at 60000fts, since at this altitude the air density is half of what it is at 40000fts ( 12000 m), drag would be cut in half, so the sustained thrust could be maintained twice as long ( 148 sec ).

Also isn't it possible to also determine at launch time the amount of fuel to accelerate to a certain speed, say Mach 4.5, depending on the speed of the aircraft?

And finaly, wouldn't it possible to determine how to use the propellent as best as possible.. If say the target is 50 km away, would it be possible to burn the propellant faster to reduce the duration of the flight.

No, it will not be possible. Solid rockets are "throttled" by molding the inner surface geometry of the solid propellant fill. The burn rate -- hence thrust and fuel consumption -- is directly proportional to the internal surface area of the propellant fill. A star shaped hollow burns faster than a hollow cylinder which burns faster than a solid cylinder lit on one end. Most boost sustain type motors have a portion of the propellant shaped with a star shaped center and the rest having a combination of cylindrical and segmented end burn grains.

In other words, you can have just about any reasonable thrust profile you want as long as you set that during the the time of manufacture. Once the motor is made, it is made. It'll burn the way it was designed and made to burn. You can't throttle it like you can a liquid fueled rocket or to a lesser degree solid fuel ramjets.

Viper`, Dwight` is correct, the only way to get a solid rocket to vary boost is to use a hybrid fuel type. A solid grain with a liquid oxidizer pressure vessel, with channels moulded into the solid fuel along which the oxidizer can flow to the combustion chamber. Through some kind of valve assembly and injector.

Not as efficient as solid fuel types but they are getting much better. HTPB/ nitrous oxide combinations seem a popular route nowadays.

Forum: F-22A Raptor

ESSM boostered AMRAAM on F-22A