Quote from: aero on 04/16/2014 06:38 amOh there is. It's equal to -g + ac where ac is centrifugal acceleration. You just have to integrate it along the trajectory you want to fly. Of course g is a function of altitude and ac is a function of horizontal velocity and altitude so the integration is a little complicated...Thanks very much to you and all the others replying on this issue. Any reference to this equation and how it can be derived?

Oh there is. It's equal to -g + ac where ac is centrifugal acceleration. You just have to integrate it along the trajectory you want to fly. Of course g is a function of altitude and ac is a function of horizontal velocity and altitude so the integration is a little complicated...

Not really, it's first principles. But you can find an explanation here.

Quote from: aero on 04/16/2014 09:17 pmNot really, it's first principles. But you can find an explanation here.Aero,I checked that page but found nothing like what I was searching for. I can easily derive the equation including the potential energy, but I read on several sources that the gravitation drag is not just something like mgh and depends on the trajectory: the more vertical you go, the more gravitational drag you pay. This is what I can't really understand.

Intuitively is the integral of g x cos(b) where b is 0 when going horizontal and Pi/2 when going vertical.

Look at the equation, GD = -g + v^2/r

That's just confusing though. Go ahead and derive the full vector formulation, starting with the vectors g and v.

On a no-atmosphere celestial body like the Moon the optimal trajectory - at least from this point of view - would be to accelerate horizontally, while remaining on the ground (no: a mass driver would feature other advantages of course), and gain altitude only when orbital speed has been achieved. At this point the maximum theoretical amount of propellant would have been used to achieve delta-v, and only the bare minimum required would be spent in potential energy while gaining altitude.

Braeunig says that the low aerodynamic drag was because the Saturn V thrust to weight ratio was low resulting in lower velocity within the lower, thicker atmosphere. He also says that the drag coefficient at max q was about 0.51. I note that Cd= 0.51 is quite low for a rocket moving just above Mach 1, which is when max q happens.

Does anyone knows where to find REAL acceleration profiles for some 1st stages other than the Saturn V one, possibly including the Falcon 9?

I've got a couple of questions...How do you calculate/estimate peak pressure in a solid rocket?

Also, if you were to scale a solid rocket motor, would the throat's surface are be proportional to the volume of the solid rocket fuel? (I basically am wondering if you wanted to scale up a solid rocket motor, what increases linearly, exponentially, etc.)

Quote from: ClaytonBirchenough on 05/04/2014 04:20 pmI've got a couple of questions...How do you calculate/estimate peak pressure in a solid rocket? Dunno.QuoteAlso, if you were to scale a solid rocket motor, would the throat's surface are be proportional to the volume of the solid rocket fuel? (I basically am wondering if you wanted to scale up a solid rocket motor, what increases linearly, exponentially, etc.)IANARS but my understanding follows.Rocket thrust is proportional to chamber pressure times throat area times thrust coefficient, the last of which is about 2 and doesn't vary much. So if you want to scale up the thrust of a solid (or liquid) rocket by a factor of X keeping chamber pressure the same you need to increase throat area by X and hence throat diameter by sqrt(X).When you increase thrust by a factor X you need to increase mass flow rate by the same factor (assuming fixed specific impulse). I believe mass flow rate in a solid, for fixed chamber pressure and propellant composition, is proportional to the area of the propellant/"air" interface in the main combustion chamber. So if you increase the thrust by a factor X increase that surface area by factor X too.Burn time, for fixed chamber pressure and propellant composition, is proportional to the (initial) thickness of the propellant in the main combustion chamber. If you don't want to change burn time when scaling your rocket you should keep this thickness unchanged - i.e. lengthen the rocket motor without changing its diameter.

How do you calculate/estimate peak pressure in a solid rocket?

Can anyone point to some info on the reentry speed capability of heatshields?What relationship is there between thickness and weight to the maximum reentry speed?If reentry speed is raised from 25,000 to (say) 27,000 mph, what effect would this have?