Ram Air Induction -ducted rocket propulsion augmentation

[bobkrech]

Fanno flow would appear to describe a loss mechanism in a supersonic flow due to to the growth of turbulant boundary layer into the main flow, dramatically increasing heat transfer to the walls. The formation of a shock wave in the main flow channel and conversion of sonic flow to subsonic flow would appear to irreversibly remove energy from the flow and IMO greatly reduce propulsive efficiency. The shock wave appears to accomplish the same function as a diffuser in a supersonic wind tunnel.

Bob

 

[retro]

Hi Bob,

Fanno flow .... I'm going to kick-the-can a bit. What if the duct wall was lined with an adiabatic heat insulator so that the stagnation temp flow was fully absorbed by the air stream? This might conserve some energy within the system. However I see a potential where as you mentioned Bob thermal or sonic choking becomes problematic. Would placing an oblique/normal shock/diffuser ahead of the constant duct help to delay supersonic issues?

TRA #279

 

[Tominator 2]

what about magnesium powder? you could probably make it a 2 stage deal. First stage getting it up to speed then a timer/altimeter would light some rocket propellant that would then light the magnesium

 

[Graham Orr]

Just as a slight correction-
Fanno flow also exists for subsonic flow.

On a particular Fanno line (on a T-s diagram), friction always has the effect of moving you to the right. This means that subsonic flow tends to move toward thermal choking and supersonic flow moves toward the same.

So, if M<1, dM/ds>0
if M>1, dM/ds<0
if M=1, dM/ds=0 or you jump to new Fanno line.

Note, however, that for M>1, you can use the intersection of the Fanno line and the Rayleigh line to find the solution for the shock. That is, when supersonic, you don't stay supersonic for long and you have shock that will move you to another solution on the subsonic branch of the Fanno line...though an increase in s occurs so energy is lost due to entropy.

 

[bobkrech]

If your interested in testing a ramjet design you might be interested in this NASA site.

https://www.grc.nasa.gov/WWW/K-12/airplane/ramjet.html

You can use their engine sim applet to test your designs. You can also download a version designed for undergraduates that allows you to change more parameters.

Ramjets operate at low compression ratios some where around 2. In the inlet the ram air is compressed, and slows down and as a result gets hot. It is this hot subsonic air that combusts with a fuel, usually in a liquid or gaseous state which adds enthalpy (heat) and mass (the fuel) to the gas flow. The net thrust you obtain is simple M-dot (mass flow per second) x delta v (feet/sec) which is the increased exhaust gas velocity that you obtain as you expand and cool the hot combustion gases in a supersonic nozzle expansion. Ths is the region of the engine where you don't want shocks as they represent non-recoverable enthalpy losses.

Bob

 

[retro]

The diffuser design has me somewhat puzzled. I want to try to get the ram mode going in the upper subsonic realm just above M.6. Searching for a book "Diffuser Design Technology" by David Japikse I think it might be helpful. However it’s pricy in the $160 range. So trying inter-library loan first. I assume the subsonic range will require a diffuser with more initial compression than an all out mach plus design would require. I ordered a NACA paper from the late forties "Theoretical Subsonic Performance of a Ducted Rocket" Hopefully that will arrive later this week and provide insights into the subsonic performance realm. I'll be quite happy if I can get M1.5 on my initial design.

 

[retro]

Would isentropic center spike design be better suited for subsonic performance?

 

[Graham Orr]

Nope. Isentropic flow is assumed...it is path independent.

 

[shockwaveriderz]

Judging by my calcs, this would have the effect of adding thrust but at the expense of lower efficiency. It'd be like an afterburner...which can prove useful especially if you have unreacted O2 in the flow.

Graham:

In your calcs did you model both the ram air induction and/or no ram air induction? And what is unreacted O2.

terry dean
barnar 16158

 

[shockwaveriderz]

The diffuser design has me somewhat puzzled. I want to try to get the ram mode going in the upper subsonic realm just above M.6. Searching for a book "Diffuser Design Technology" by David Japikse I think it might be helpful. However it’s pricy in the $160 range. So trying inter-library loan first. I assume the subsonic range will require a diffuser with more initial compression than an all out mach plus design would require. I ordered a NACA paper from the late forties "Theoretical Subsonic Performance of a Ducted Rocket" Hopefully that will arrive later this week and provide insights into the subsonic performance realm. I'll be quite happy if I can get M1.5 on my initial design.

Retro:

what kind of avearge Mach numbers would I see with typical ABCDE Estes BP motors? 0.2 to 0.4?

terry dean
barnar 16158

 

[retro]

what kind of avearge Mach numbers would I see with typical ABCDE Estes BP motors? 0.2 to 0.4

To many variables involved to answer your question without detailed dimensional, material, design, data it is impossible to know. I saw a NACA paper a few days ago based very small diameters test in the 1 inch range. I suspect Reynolds’s number and boundary layer effects would be more dramatic at that small of scale

And what is unreacted O2

I suspect Graham was referring to unburnt O2 (oxygen) from the ejector (booster) motor exhaust, entrained free stream O2 or both flowing into the combustion chamber.

Got the Diffusser book today (library loan... Freebeeee) looks informative already cleared up a few things just skimming through it.

 

[Graham Orr]

Retro:

what kind of avearge Mach numbers would I see with typical ABCDE Estes BP motors? 0.2 to 0.4?

terry dean
barnar 16158

Although it isn't by any means perfect (as a CD nozzle) you can expect exhaust velocities of about 750m/sec (2500 ft/sec) out of a BP rocket motor. If the exhaust is at about 800K this results in a mach number of about 1.4 relative to the flow conditions. Alternatively, relative to ambient conditions it is about mach 2.2. (Temperature dictates the speed of sound leading to higher speed of sound in the flow and lower mach number in the flow...even though it is a specific velocity)

Retro,
You are correct. I was referring to unreacted (non-reduced) oxygen in the flow. However, if you are working with an electrothermal system, all of the flow is unreacted... Which means you can add even more heat via combustion and jump to a new Rayleigh line. This, however, decreases performance at the cost of thrust.

 

[shockwaveriderz]

Their was a OOP kit called the Ram-Jet.

I saw a 4" shorta version of that in the late 80s with the motor moved up higher into the body tube so it would mix with the air being ramed into the 4" ram tube.

IE a 4" body tube has a 2.56" body tube and nose cone inside it up top and tapers down to 38mm motor mount tube that stop short near the rear.

The fake internal static tubine fins hold the internal tubes to the 4" body/ram tube.

It makes a nice Kewl looking smoke trail as the Krushnic effect mixes with the air stream.

Art:

Do you remem who offered this 4 sale?


terry dean
nar 16158

 

[shockwaveriderz]

. I ordered a NACA paper from the late forties "Theoretical Subsonic Performance of a Ducted Rocket" Hopefully that will arrive later this week and provide insights into the subsonic performance realm. .

Retro:
When I 1st searched for this it was for sale; last night I accidently found it free online here:

https://hdl.handle.net/2060/19630043472

IT is VERY GOOD but mathematically HEAVY!

terry dean
nar 16158

 

[bobkrech]

Retro:

what kind of avearge Mach numbers would I see with typical ABCDE Estes BP motors? 0.2 to 0.4?

terry dean
barnar 16158

The exhaust velocity of any rocket is by definition supersonic. The exhaust velocity V = g*Isp where g = 9.8 m/s^2 gravitation acceleration and Isp is the specific impulse in seconds.

BP motors typical have a specific impulse between ~70 to ~80 sec so the exhaust velocity will typically range from ~690 ms/ to ~790 m/s.

You can determine the exhaust velocity for a certified rocket motor for the certification data sheet.

V (m/s) = total impulse (N-s) / propellant mas (kg)

For example consider the Estes B6 motors https://www.nar.org/SandT/pdf/Estes/B6.pdf

The total impulse is 4.33 newton-seconds and the propellant mass is 5.6 grams

V m/s = 4.33 N-s / (5.6 g/1000 g/kg) = 733 m/s

The specific impulse, Isp = 733 m/s / 9.8 m/s^2) = 78.9 s.

Bob

 

[Graham Orr]

Bob,
Isn't Isp/(m*g) yield the equivalent velocity (Veq)? I thought you would take this and divide by your Cf to get exhaust velocity (Ve).

F=mdot*Ve+At&#8710;P = mdot*Veq = mdot*Ve*Cf

Cf for estes motor is probably around 1.3 or so...

F/mdot=Veq

Perhaps I am not remembering this correctly...

Either way, as you state, the flow is going to be supersonic and as a very rough estimate around mach 1.5 relative to the exit gas temperature.

EDIT: The above math is incorrect. I meant to say Isp =I/(mg). My bad. See posts below...

 

[Graham Orr]

Retro:
When I 1st searched for this it was for sale; last night I accidently found it free online here:

https://hdl.handle.net/2060/19630043472

IT is VERY GOOD but mathematically HEAVY!

terry dean
nar 16158

I hope it is mathematically heavy... considering you probably need to work hard to get the conditions where the effect would help you. Typically in compressible flow, you have to work through the math to get the conditions at which the flow performs. Because of the nature of compressible flow, a lot of counterintuitive things occur and thus you can't just rely on your everyday perception of fluid mechanics but really have to form a trust in the thermodynamics of the flow. After awhile, you can start to get a bit of intuition as to what super/hypersonic fluids are going to do...

For instance: Most people think that a sharp tipped nosecone (with the leading edge less than the mach angle) is best for supersonic aerodynamics. This is only the case in a very small region of supersonic flow...and what you'll find out is that you actually want to induce a detached shockwave from the nose at higher mach numbers...which will aid in overall drag reduction and decreased aerodynamic heating. This can be done with a blunt-tipped nose but how many of those do you see in high performance amateur built rockets? Not many...but you definitely see it in NASA-style rockets.

But of course, there is always the consideration as to how long the object is supersonic verses subsonic and blunt bodies are not great at subsonic...so maybe secretly everyone who flies high mach number rockets secretly do some calculations without telling anyone...

 

[bobkrech]

Bob,
Isn't Isp/(m*g) yield the equivalent velocity (Veq)? I thought you would take this and divide by your Cf to get exhaust velocity (Ve).

F=mdot*Ve+At&#8710;P = mdot*Veq = mdot*Ve*Cf

Cf for estes motor is probably around 1.3 or so...

F/mdot=Veq

Perhaps I am not remembering this correctly...

Either way, as you state, the flow is going to be supersonic and as a very rough estimate around mach 1.5 relative to the exit gas temperature.

I may have confused you by stating propellant weight instead of propellant mass which I have corrected in editing.

For the student:

Show that F(N) /mdot (kg/s) = Veq (m/s) = TI (N-s) / m (kg)

hint: multiply LHS by 1 = 1s/1s

Exhaust velocity expressed in Mach number is conventionally referred to ambient temperature, not the hot gas temperature.

Bob

 

[Graham Orr]

Oops. I made a mistake too.

I meant to say:

Isp = I/(m*g)

Then, Veq= Isp*g = I/m

I think we're on the same page now

 

[retro]

Retro:
When I 1st searched for this it was for sale; last night I accidently found it free online here:

Bummer.... I purchased my copy from the NASA server. Guess I should have spent more time searching. Yes JPL #N63-86365 is a good resource, I'm still digesting my copy. Each time I read it I gleen more information and understanding. Here are two other resource I have found to be very informative The book was under used for under$10 at amazon.com.

NACA RM L54G21a
Available at Aerade
https://aerade.cranfield.ac.uk/results.php?Simplequery=inlet&page=14&sf=KW&st=AND

The book is dated however it is an excellent diffuser design resource! This book assumes the reader has little previous diffuser knowledge. And it lays out all the basics. Choking, inlet start, swallowing, subsonic to supersonic section by section, including compressible and incompressible flow dynamics and calculations, and the most complete listing of variables and symbols I've ever seen

"Supersonic Inlet Diffusers and Introduction To Internal Aerodynamics"
by Rudolf Hermann
1956

 

[Graham Orr]

Okay, so I have a revision on my previous design:
This one is a shock-Raleigh engine. It has a strong shock (or normal) in the inlet and heat transfer (via combustion) in the constant area portion. Finally, the flow is diverged.

Note: The unmarked peach line in the second pic is a Fanno line.

Note Again: The Isp graph is a little shocking. The engine cannot operate at those SUPER high Isps due to the air/fuel ratio. Also, I think there is a bug in the simulation somewhere...so don't ream me for that Isp curve. i expected something around 2000-3000 sec but other than the fact that the engine doesn't produce nearly any thrust in the super high Isp range I'd have to say even if is possible the thrust level is too low to counter drag and such. I'll work this out before next monday.

 

[retro]

Neat concept....with those temps you may need a more exotic fuel such as LH2 or triple point. LH2 vaporized by the heat and injected through a high velocity nozzle can deliver 5,000 ISP alone. You also release the stored energy from the process utilized to cool the fuel. If you are a "hands on" person and have a true passion for this kind of stuff I have a project in the works that might interest you.....????

Retro

Vis Viva

 

[powderburner]

The total impulse is 4.33 newton-seconds and the propellant mass is 5.6 grams

V m/s = 4.33 N-s / (5.6 g/1000 g/kg) = 733 m/s

The specific impulse, Isp = 733 m/s / 9.8 m/s^2) = 78.9 s.

Bob,

If I read these notes correctly, these numbers refer to an overall or "average" mass flow condition.

Any hints on how to model an initial thrust peak separately from the "sustaining" portion of a typical BP motor burn? (so that net impulse still comes out right) (or does this go too deeply into complex math to try to present here in simple form?)

I imagine that all those transient conditions will be a booger to analyze...

 

[Graham Orr]

Neat concept....with those temps you may need a more exotic fuel such as LH2 or triple point. LH2 vaporized by the heat and injected through a high velocity nozzle can deliver 5,000 ISP alone. You also release the stored energy from the process utilized to cool the fuel. If you are a "hands on" person and have a true passion for this kind of stuff I have a project in the works that might interest you.....????

Retro

Vis Viva

Hey Retro,
Either that, or a microwave emitter powerful enough to vaporize the water in Gotham City....errr....I mean a microwave emitter tuned to the N2 triple bond resonant frequency. In the next 20 years we may see RF devices efficient enough to do this so such an engine could be electrically powered.

The Isp quotes was by using the energy density of gasoline... you're right in that gas tops out at about 3700K in adiabatic flame temp (I believe).

Oh, I got your email/message. I'm thinking about it...got a lot of proposals and whitepapers on the board right now...though there is a possibility that one of my potential projects could provide facilities for testing for your project.

 

[Graham Orr]

Bob,

If I read these notes correctly, these numbers refer to an overall or "average" mass flow condition.

Any hints on how to model an initial thrust peak separately from the "sustaining" portion of a typical BP motor burn? (so that net impulse still comes out right) (or does this go too deeply into complex math to try to present here in simple form?)

I imagine that all those transient conditions will be a booger to analyze...

If you can measure the mass flux at each point in time in the motor then you would be able to extract instantaneous Isp. Would be:

Isp=F/(g*mdot)

You can make approximations as to what mdot is by ideal gas law and idealized flow through the nozzle. If you had a method of measuring pressure and temperature during the burn you could get a nozzle flow-rate out of that. This could give you instantaneous Isp. As you can imagine, however, acquiring instantaneous P and T are quite challenging.

From a conceptual standpoint, it would make sense if your Isp dropped during the sustained thrust period due to the lower operating pressure of the motor (with same nozzle geometry, roughly).

 

[bobkrech]

Bob,

If I read these notes correctly, these numbers refer to an overall or "average" mass flow condition.

Any hints on how to model an initial thrust peak separately from the "sustaining" portion of a typical BP motor burn? (so that net impulse still comes out right) (or does this go too deeply into complex math to try to present here in simple form?)

I imagine that all those transient conditions will be a booger to analyze...

It can be done but it's really not very important for this application.

I don't have a clue what chamber pressures exist in Estes motors. From Isp calculations and thrust measurements, I'd guess it's around 100 psi or so, but that really seems low, however since the nozzles are pressed in clay and the casings are paper, that might be all there is. If you could get up in the 150 to 500 psi range, the Isp would be significantly higher.

For every chamber pressure in a BP motor there is an optimum expansion ratio which can range from about 2 to 12 for chamber pressures ranging from 10 to 68 atm. You can only design the expansion nozzle to provide the maximum Isp at one chamber pressure so it will be sub optimal for all others. If you look at the thrust curves for various Estes BP motors, you will see that the peak thrust to sustaining thrust ratios varies ~2.5:1 to 4:1. If I were designing the motor I would design for max Isp during the sustainer part of the burn and ot worry about Isp during the initial thrust spike, as over expansion is less of a hit on Isp than underexpansion.

It's not clear exactly what optimization has been done on most BP motors as the Isp appears to be lower than what I would expect. You can play around with PROPEP and determine what the optimum Isp should be for a given chamber pressure. An option will allow you to determine the optimum expansion ratio if yu can get an idea of what the chamber pressure is.

Bob

 

[shockwaveriderz]

you basically have two chamber pressures in the typical Estes BP motor. During the approximate .2 sec thrust spike, at max/peak thrust you have the maximum chamber pressure which I would estimate to be 150-175 psi; while the sustainer chamber presure is usually around 75 psi; this of course is with the approximate 2.0 expansion ratio that the Estes nozzles have.

edit

I forgot to say that the values above are for an C6 motor. The chamber pressures that exist in the thrust spike/sustainer values of other Estes motors is in that ballpark ie.... the chamber pressures are generally higher for the DE motors while generally lower for smaller motors.

terry dean
nar 16158

 

[retro]

Given the degree of scrutiny, legal fees and never ending regulations this hobby has endured over the last few years I think its time to look for alternate fuels, oxidizers and propulsion for rockets that require more than 62.5 grams of AP based propellant. Augmentation and ducted rocket research is a step in that direction.

 

[bobkrech]

Given the degree of scrutiny, legal fees and never ending regulations this hobby has endured over the last few years I think its time to look for alternate fuels, oxidizers and propulsion for rockets that require more than 62.5 grams of AP based propellant. Augmentation and ducted rocket research is a step in that direction.

Retro

A ram jet is not a rocket. Rocket do not need air to operate. Jets do.

Bob

 

[retro]

Bob you are right about the ramjet it’s not a rocket. However a ducted rocket does utilize entrained air. Now when a ducted rocket is a rocket or a ram rocket or an ejector rocket or a ramjet seems kind of blurry? In the case of a RBCC it depends on which part of the cycle it’s operating in at the moment. Ramjet, ram rocket, ducted rocket, ejector rocket, air rocket, augmented rocket, rich, statometric or lean ratio all I really care is that it goes upstairs like a-bat-out-of- h---.

The concept I'm looking at could even be two stage (Lil Augie II). First stage is pure rocket boost, then that stage drops out of the bottom of the hollow tube, second fires at the top of the tube and it now becomes an ejector rocket for a few seconds igniting the hollow fuel core liner below. As the second stage burns out the ejector rocket is going close to mach 1 and transitions to solid fuel (paraffin) ramjet mode consuming the remaining solid fuel liner inside the hollow tube.