Experimental validation of rotating detonation for rocket propulsion

Abstract

Space travel requires high-powered, efficient rocket propulsion systems for controllable launch vehicles and safe planetary entry. Interplanetary travel will rely on energy-dense propellants to produce thrust via combustion as the heat generation process to convert chemical to thermal energy. In propulsion devices, combustion can occur through deflagration or detonation, each having vastly different characteristics. Deflagration is subsonic burning at effectively constant pressure and is the main means of thermal energy generation in modern rockets. Alternatively, detonation is a supersonic combustion-driven shock offering several advantages. Detonations entail compact heat release zones at elevated local pressure and temperature. Specifically, rotating detonation rocket engines (RDREs) use detonation as the primary means of energy conversion, producing more useful available work compared to equivalent deflagration-based devices; detonation-based combustion is poised to radically improve rocket performance compared to today's constant pressure engines, producing up to 10[Formula: see text] increased thrust. This new propulsion cycle will also reduce thruster size and/or weight, lower injection pressures, and are less susceptible to engine-damaging acoustic instabilities. Here we present a collective effort to benchmark performance and standardize operability of rotating detonation rocket engines to develop the RDRE technology readiness level towards a flight demonstration. Key detonation physics unique to RDREs, driving consistency and control of chamber dynamics across the engine operating envelope, are identified and addressed to drive down the variability and stochasticity observed in previous studies. This effort demonstrates an RDRE operating consistently across multiple facilities, validating this technology's performance as the foundation of RDRE architecture for future aerospace applications.

Figure 1

Views of a rotating detonation rocket engine including the (a) hardware schematic, and (b) engine firing of the canonical hardware.

Figure 2

Engine performance summary showing (a) thrust F and (b) specific impulse $I_s$ as a function of total propellant mass flow rate ${\dot{m}}_{\text{tot}}$ at $\varphi$ = 1.1. For certain tests, error bars are smaller than the symbols shown.

Figure 3

Flow condition test matrix for the canonical RDRE hardware tested at the four research facilities. For certain tests, error bars are smaller than the symbols shown.

Figure 4

Summary of the detonation mode characteristics as a function of total propellant mass flow rate at $\varphi$ = 1.1, depicting the (a) number of waves m, and (b) average wave speed $U_{\text{wv}}$. For certain tests, error bars are smaller than the symbols shown.

Figure 5

Injection pressure drop of the (a) oxidizer and (b) fuel as a function of total propellant mass flow rate at $\varphi$ = 1.1.

Figure 6

Axial chamber pressure profiles for (a) $\varphi$ = 1.1 at ${\dot{m}}_{\text{tot}}$ = 0.272 kg/s, (b) $\varphi$ = 1.1 at ${\dot{m}}_{\text{tot}}$ = 0.363 kg/s and (c) $\varphi$ = 1.7 at ${\dot{m}}_{\text{tot}}$ = 0.272 kg/s, with (d) locations of CTAP pressure sensors indicated on schematic of laboratory scale rotating detonation rocket engine.

Figure 7

Transient run data for the $\varphi$ = 1.1, ${\dot{m}}_{\text{tot}}$ = 0.272 kg/s (nominal) condition showing (a) thrust, (b) CTAP1 sensor chamber pressure, (c) oxidizer and (d) fuel plenum pressures.