Although a previous answer has been selected I think the asker needs a bit more info to understand what sort of questions you should be asking about a rocket. In particular, the Tsiolkovsky rocket equation is only ever used to compute a required mass ratio or delta-V. It's essentially a staging tool; given some total delta-V what are the number of burns and stages I need at what MR? What number of burns or stages would best fit the problem? Is this possible given constraints on payload and inert masses? It should not be used for any sort of real flight dynamics including computing vehicle acceleration, which in an of itself is not typically something people worry too much about. There are some simple models for flight dynamics to consider things like g-t losses or basic aerodynamic effects, but when baselining a rocket you would incorporate all of those into a rough efficiency--say you will have a 10-20% penalty in propellant mass.
The basic rocket relationships which hold regardless of the actual model of the rocket flow are:
(1) Tsiolkovsky equation: $MR = e^{\Delta V/c}$
where $c$ is the effective exhaust velocity, viz. the velocity of the propellant gases after pressure losses have been taken into account. This is used for staging, delta-V, and mass ratio calculations.
(2) Basic Thrust Equation: $F = \dot{m}c$.
This equation is used to determine what the thrust the rocket will produce is given $c$, since most ways of analyzing a rocket return $c$ rather than $F$ (thrust). Thrust is more important to consider in whether a rocket is capable of executing a certain maneuver rather than whether the rocket itself is a good design. High thrust buy low Isp is almost always worse than slightly lower thrust at higher Isp.
(3) Constant Burn Rate Assumption: $m_b = \dot{m}\Delta t$
where $m_b$ is the mass of the burned propellant. This is less of a fundamental result than an assumption. We assume a constant burn rate most of the time, since this is easy to work with from an engineering perspective.
You mention specific impulse, which is formally defined as thrust per unit weight of mass expended, but as you can see from (2) this isn't actually a different parameter from the effective exhaust velocity, e.g.
(4) $I_{sp} = c/g$
Where $g= 9.81$ m/s$^2$, viz. the acceleration near earth, and is only in there because of the definition of $I_{sp}$.
Contrary to the previous answer it is not enough to know simply what your propellants are to determine $c$ or $I_{sp}$--then rocketry would be stupid easy! In general you need the type of rocket (e.g. solid, liquid monopropellant, liquid bipropellant, etc.), type of propellants, mixture ratio of the propellants, nozzle geometry (especially the expansion and contraction ratios), and thermochemical states of the propellants in the combustion chamber (which leads into engine cycles, injector theory, chemical kinetics of the propellant flow and a whole host of other topics). This doesn't even begin to touch on loss mechanisms for a non-ideal rocket, most of which affect $c$ and $I_{sp}$ or whether or not real rocket could work for a given design (e.g. cooling methods, structural integrity, combustion stability). A good place to start if you want to do a first order estimate of $c$ is by selecting some propellants, running a thermochemical code, like NASA's CEA code, and use the results in an isentropic analysis (see a book like Sutton, Rocket Propulsion Elements for more detail). This gives you first order estimates of $c$, and even $\dot{m}$ if you learn how to use $c^*$ velocities correctly.
