Influence of a Conical Axial Injector on Hybrid Rocket Performance

Abstract

This paper analyzes the results obtained from a series of static firings of a lab-scaled hybrid rocket in which gaseous oxygen was supplied into axial-symmetric polyethylene cylindrical grains through two different injector configurations: an axial conical subsonic nozzle and a radial injector. The axial injector is interesting because of its relatively easy design, the higher regression rates, and the noticeably stable motor operation. To exploit its qualities, not only the assessment of the regression rate but rather the entire behavior of the motor is required. For the investigated set of operating conditions, the instantaneous regression rates exhibit a time dependence caused by the impinging jet zone dynamics, while the average regression rates are higher and less mass flux dependent than those achieved with the radial injection motor and expected from the classical turbulent-boundary-layer diffusionlimited theory. A comparison to the data from the radial injector was further drawn in terms of combustion efficiency and fuel regression uniformity. Concerning combustion stability, some observations are made. The radial injector, at the same mass flux and pressure, produces lower regression rates, high pressure oscillations, and worse combustion efficiency, but more uniform fuel consumption. Nomenclatureof ultrasounds in reference conditions c * = characteristic exhaust velocity c * 0 = theoretical characteristic exhaust velocity calculated at O/F c * = theoretical characteristic exhaust velocity calculated at OF D = space average port diameter D 0 = port initial diameter D 2 = average port final diameter D 2x = local port final diameter d = diameter E a = activation energy F = motor thrust G = mass flux H v = effective heat of vaporization I m = mixing index L = grain length L c = prechamber length L * = chamber characteristic length l = exponent of the Reynolds number in the mixing index definition m = exponent of the geometrical ratio in the mixing index definitioṅ m f = fuel mass flow ratė m ox = oxidizer mass flow rate = isentropic one-dimensional mass flow rate through the nozzle N = number of data points in the port diameter profiles n = exponent of mass flux OF = spatially averaged oxidizer-to-fuel ratio OF = mean value of spatially averaged oxidizer-to-fuel ratios O/F = average oxidizer to average fuel mass ratio p a = ambient pressure p e = nozzle-exit pressure R = gas constant Re = Reynolds numbeṙ r = space average regression ratė r x = local regression rate s = fuel grain thickness T w = fuel surface temperature t = time V e = isentropic one-dimensional velocity at nozzle exit x = axial abscissa z = nondimensional axial abscissa (x/L) α = nozzle divergence cone half-angle M f = measured solid fuel mass loss M f = computed solid fuel mass loss t = time step t o = ultrasounds time of flight η = c * efficiency, ratio between the actual and theoretical c * th η 0 = c * efficiency calculated at O/F λ = momentum-thrust reduction coefficient μ = gas viscosity ρ f = solid fuel density σ = standard deviation...

Estimation of Hybrid Rocket Nozzle Throat Erosion History

“…The large fluctuation in the initial 2 s is due to the start-up unsteadiness and should be ignored. 7) After this period to 4 s, the nozzle throat area is nearly constant, showing that the nozzle throat erosion started 4 s after the startup. This result agrees with Fig.…”

“…Carmicino et al showed experimentally that the efficiency η is almost constant during firing except for the initial and final transients unless the flow field structure greatly changes. 7) Figure 12 shows histories of η with and without considering nozzle throat erosion by gray and black solid lines, respectively. The efficiency η without considering the nozzle erosion begins decreasing at 4 s. On other hand, the history with considering the nozzle erosion is more reasonable, being almost constant.…”

“…Some researchers have developed various reconstruction techniques [5][6][7] to obtain ξ from measurable data such as chamber pressure P c or oxidizer flow rate o m  . Common to all of the reconstruction techniques ever proposed, nozzle throat area A t is necessary as an input data.…”

Experimental Study of Fuel Regression Rate in a HTPB/N2O Hybrid Rocket Motor

“…Many theoretical and experimental studies have shown that the fuel regression rate in a hybrid motor is highly dependent on the mass ux passing through the fuel port [21][22][23]. Thus, in most studies, the regression rate only correlates with a parameter given by: _ r = a G n o : (9) For a xed amount of fuel and oxidizer, it has been reported that the fuel regression weakly depends on the pressure of the combustion chamber, longitudinal distance from the edge of the fuel, and port diameter of the fuel [24][25].…”

Method for Determining Nozzle-Throat-Erosion History in Hybrid Rockets

“…NTRT builds from the governing equations and concepts of the reconstruction techniques introduced in Ref. [9][10][11] to reconstruct the nozzle throat radius R t and ξ histories through data reduction of commonly measured experimental values: (1) oxidizer mass flowrate, ox m  ; (2) (aft) chamber pressure, P c ; (3) thrust, F ; (4) overall fuel mass consumed, M f ; and either (5a) final nozzle throat radius, R t,f ; or (5b) nozzle exit pressure, P e , in the case that R t,f is not obtainable.…”