Optimization of a spacecraft's interplanetary trajectory and electric propulsion system remains a complex and difficult problem. Simultaneously solving for the optimal trajectory, power level, and exhaust velocity can be difficult and time consuming. If the power system's technology level is unknown, multiple optimizations must be conducted to map out the trade space. Trajectories with constant-power, solar-power, variable-specific-impulse, and constant-specific-impulse low-thrust propulsion systems are analyzed and optimized. The technological variables, power system specific mass, propellant tank coefficient, structural coefficient, and the launch vehicle are integrated into the cost function allowing for maximization of the payload mass fraction. A classical solution is reviewed that allows trade studies to be conducted for constant-power, variable exhaust velocity systems. The analysis is expanded to include bounded-power constant specific impulse systems and solar electric propulsion spacecraft with constant and variable exhaust velocity engines. The cost function and mass fractions are dimensionless to allow for scaling of the spacecraft systems. Nomenclature a = thrust acceleration, m/s 2 a 0 = nominal acceleration spacecraft can deliver at launch, m/s 2 a * = characteristic acceleration, m/s 2 C 3 = launch energy, km/s 2 c = exhaust velocity, m/s c * = optimal exhaust velocity, m/s f (C 3 ) = launch-vehicle mass fraction G = gravitational vector, m/s 2 g = gravity constant defined as 9.8 m/s 2 I sp = specific impulse, s J 1 = payload mass fraction J 2 = quadratic cost function for constant-power variable I sp k 1 = control variable that determines the power fraction utilized by the spacecraft k 2 = path and time-dependent function that determines the maximum power fraction as a function of the power level at launch m = spacecraft mass, kg m prop = propellant mass, kg m ps = mass of power system, kg m * prop = propellant mass fractioṅ m = mass flow rate, kg/s P j = jet power, W P ppu = power into power processing unit, W r = position vector, m T = thrust, N t = time, s t burn = engine on time, s Z = power weighted burn time, ṡ Z = current-power-to-initial-power utilization ratio . Associate Fellow AIAA. α = power system's specific mass, kg/W β = jet-power-to-spacecraft-mass ratio, W/kg β * 0 = optimal initial jet-power-to-mass ratio, W/kg = power mass fraction of spacecraft V = velocity change imparted by engine, m/s V * = normalized V ζ = launch mass utilization factor η = propellant tank coefficient η engine = efficiency of electric propulsion engine η ppu = efficiency of power processing unit = structural coefficient τ = integration dummy variable φ = tank and structural coefficient scaling factor ψ = propellant and tank scaling factorSubscripts and Superscripts f = final state 0 = initial state · = time derivative