I. O. Shamshin

Energy efficiency of a continuous-detonation combustion chamber

Systematic experimental and computational studies of the energy efficiency of continuous-detonation combustors (CDCs) have been performed. A small-size and a large-size CDCs using hydrogen as fuel and oxygen or air as oxidizer have been developed and tested. It was first experimentally proved that the Zel’dovich thermodynamic cycle with continuous-detonation combustion of a hydrogen-oxygen mixture in an annular combustor is more efficient than the Brayton thermodynamic cycle with continuous combustion of the mixture, other things being equal. The specific impulse of a small-size bench-scale rocket engine with a 50 mm diameter CDC operating in the continuous-detonation mode was 6–7% higher than that in the continuous combustion mode of operation. The measured fuel-based specific impulse for the large-size CDC of 406 mm diameter running on a hydrogen-air mixture was at a level of 3000 s. Three-dimensional calculations to optimize the structure and operation mode of the large-size CDC have shown that when running on a combustible mixture with a nearly stoichiometric overall composition, the specific impulse can be increased to ≈4200 s.

Experimental proof of the energy efficiency of the Zel’dovich thermodynamic cycle

207 In 1940, Zel’dovich advanced the idea of the possi bility of using the energy of detonation combustion [1]. According to his estimates, the thermodynamic efficiency of a cycle with detonation combustion of fuel can significantly exceed the efficiency of the Bray ton cycle with constant pressure combustion. Later, these theoretical conclusions and estimates [1] were confirmed by thermodynamic calculations and multi dimensional gas dynamics calculations with consid eration of dissipative processes. For example, thermo dynamic calculations [2] showed that the efficiency of the Zel’dovich cycle (as the detonation combustion cycle is called now) can be 20–30% higher than that of the Brayton cycle, and multidimensional gas dynam ics calculations of the process in a liquid propellant rocket engine with detonation combustion [3] pre dicted a 13–15% gain in comparison with a usual liq uid propellant rocket engine. Although the theoretical conclusions about the energy efficiency of the Zel’dovich cycle are unquestioned, there have been no direct experimental proofs of these conclusions to date. In this work, for the first time, we experimentally proved that the Zel’dovich thermodynamic cycle with continuous detonation combustion of a hydrogen– oxygen mixture in an annular chamber is more effi cient than the Brayton thermodynamic cycle with continuous combustion of the same mixture, other conditions being equal. The specific impulse of a rocket engine prototype operating in the continuous detonation mode was 6–7% higher than that in the continuous combustion mode. There are two main designs to perform detonation combustion: in detonation waves circulating continu ously in the tangential direction across an annular combustion chamber (continuous detonation cham bers [4]) and in periodic detonation waves travelling along a combustion chamber (pulse detonation cham bers [5]). At the present time, the problem of experi mental implementation of continuous and pulse deto nation combustion of various fuels (from hydrogen to aviation kerosene) with various oxidizers (air, oxygen enriched air, oxygen) has generally been solved. The issue is that there have still been no experimental proofs of the energy efficiency of the Zel’dovich cycle. For example, all the known experiments on continu ous detonation combustion in annular combustion chambers were accompanied by high pressure loss in propellant component feed systems and demonstrated low process efficiency below the ideal efficiency of continuous combustion chambers. No direct compar ison has been made between the measured thrust per formances of pulse detonation and pulse combustion jet engines. The purpose of this work was to experimentally prove the energy efficiency of the Zel’dovich cycle by directly comparing the thrust performances of a rocket engine prototype operating in the continuous detona tion and continuous combustion modes. For this pur pose, we designed and built a test stand and a rocket engine prototype, which can operate in the modes of continuous detonation and continuous combustion of a mixture of gaseous hydrogen and gaseous oxygen. The test stand consists of hydrogen and oxygen receivers 0.64 and 0.32 m3 in volume, respectively, high performance system of fast response valves, large section propellant component feed manifolds, thrust table, and precision system for measuring the thrust and the propellant component feed pressures. The maximal mass flow rate of the propellant mixture that can be reached on this stand is close to 1.5 kg/s. A rocket engine prototype is an annular combus tion chamber to which, to one end face, an injector PHYSICAL CHEMISTRY

Propagation of shock and detonation waves in channels with U-shaped bends of limiting curvature

Systematic experimental and theoretical studies of the propagation of shock and detonation waves in cylindrical tubes and planar channels with two U-shaped bends of limiting curvature were performed. It was demonstrated that U-shaped bends substantially facilitate detonation initiation in gases. The minimum shock wave velocity required to initiate the detonation of a stoichiometric propane-air mixture under normal conditions in a near-critical diameter tube with two U-shaped bends of limiting curvature was found to be ∼800 m/s.

Demonstrator of continuous-detonation air-breathing ramjet: Wind tunnel data

First experimental investigations were carried out into the detonation combustion of hydrogen in a demonstrator of an original-design air-breathing ramjet while blowing with an air flow at Mach 4 to 8 in an impulse wind tunnel, and for the first time under these conditions, continuous spin and longitudinal pulsed modes of detonation combustion of hydrogen in an annular combustor were detected.

Thrust characteristics of a pulse detonation engine operating on a liquid hydrocarbon fuel

A demonstrator of a pulse detonation combustion chamber of original design based on a cyclic deflagration- to-detonation transition in a mixture of separately fed liquid hydrocarbon fuel (propane–butane mixture) and air was developed. Fire tests of the demonstrator with an attached air duct, operating frequencies of up to 20 Hz, were performed on a thrust measurement bench. During the tests, wave processes in the gasdynamic duct were monitored and fuel consumption rate and thrust force were measured. At a frequency of operation of the demonstrator within 2–15 Hz, the fuel-based specific impulse was ~1000 s. It is shown that a partial filling of the gasdynamic duct with fuel mixture makes it possible to increase the specific impulse up to ~1100 s.

Momentum transfer from a shock wave to a bubbly liquid

The transfer of momentum from shock waves of various intensities (from 0.05 to 0.5 MPa) to a water column containing air bubbles of a mean diameter of 2.5 mm is studied both experimentally and by numerical simulation. The experiments are performed in a vertical hydrodynamic shock tube with a rectangular cross section of 50 × 100 mm and a length of 1980 mm. The tube consists of a 495-mm-long high-pressure section, 495-mm-long low-pressure section, and 990-mm-long test section filled with water and equipped with a bubble generator. Experiments have demonstrated that, as the gas content in the water increases from 0 to 30 vol %, the momentum transferred from the shock wave to the bubbly water increases smoothly, leveling off at a volumetric gas content of 20–25%. The experimental and 2D-simulation dependences of the shock wave velocity and the velocity of the bubbly liquid behind the shock wave front on the volumetric gas content are in close agreement.

Initiation of detonation in a tube with a profiled central body

Numerical simulation of shock wave propagation in a tube with a central body proved the possibility of significant facilitation of the deflagrationtodetona� tion transition in a methane-air mixture. It was shown that the central body should meet some requirements: it should block approximately 60% of the tube cross section, and should have a parabolic forebody profile with an angle of attack of ~40° and a parabolic or ellip� tic afterbody profile with an angle of convergence of no more than 7°. In this case, detonation can be initi� ated by a shock wave with a Mach number of 3.5 over a length of ~0.5 m for a time of ~0.2 ms.

Deflagration-to-detonation transition in the gas–liquid-fuel film system

A deflagration-to-detonation transition was experimentally detected for the first time in a channel with a thin wall liquid-fuel film and a gaseous oxidizer using a weak ignition source, which generates no primary shock wave of any significant intensity. In a number of tests, a low-velocity quasi-stationary detonationlike combustion front traveling at an average velocity of 700–900 m/s was recorded; the structure of this front included a leading shock wave and a reaction zone following after a time delay of 80 to 150 μs.

Magnetohydrodynamic effects of heterogeneous spray detonation

Electrical power on the board of an aircraft with a liquid-fuel pulse-detonation engine is proposed to be produced by a magnetohydrodynamic (MHD) generator placed at the outlet of the jet nozzle. MHD effects of the pulsed heterogeneous (spray) detonation of n-heptane–oxygen mixtures with addition of a potassium carbonate aqueous solution (ionizing additive) to the detonation products are studied. The MHD channel electrodes are demonstrated to steadily generate voltage pulses with an amplitude of up to 3 V. Three types of shape of individual pulses are observed: single pulses, double pulses, and double pulses with a shortterm change in the sign of the voltage. Increasing the frequency of operation of the PDE from 20 to 40 Hz does not affect the amplitude and shape of the voltage pulse. Reducing the magnetic induction from 0.6 to 0.3 T decreases the amplitude of the voltage pulse across the MHD channel electrodes, but the shape of individual pulses remains virtually unchanged. Leaning the fuel mixture reduces the generated voltage. Addition of mechanoactivated Mg-MoO3 energetic nanocomposite to the fuel mixture does not cause significant changes in the shape and amplitude of voltage pulses.

Calculation of shock wave propagation in water containing reactive gas bubbles

The entry of a shock wave from air into water containing reactive gas (stoichiometric acetylene–oxygen mixture) bubbles uniformly distributed over the volume of the liquid has been numerically investigated using equations describing two-phase compressible viscous reactive flow. It has been demonstrated that a steady-state supersonic self-sustaining reaction front with rapid and complete fuel burnout in the leading shock wave can propagate in this bubbly medium. This reaction front can be treated as a detonation-like front or “bubble detonation.” The calculated and measured velocities of the bubble detonation wave have been compared at initial gas volume fraction of 2 to 6%. The observed and calculated data are in satisfactory qualitative and quantitative agreement. The structure of the bubble detonation wave has been numerically studied. In this wave, the gas volume fraction behind the leading front is approximately 3–4 times higher than in the pressure wave that propagates in water with air bubbles when the other initial conditions are the same. The bubble detonation wave can form after the penetration of the shock wave to a small depth (~300 mm) into the column of the bubbly medium. The model suggested here can be used to find optimum conditions for maximizing the efficiency of momentum transfer from the pressure wave to the bubbly medium in promising hydrojet pulse detonation engines.