Aerospike Nozzle for Rotating Detonation Engine Application is sponsored by NASA. Graduate research on improving the efficiency of rotating detonation engines by using aerospike nozzle technologies.

This proposal presents a graduate MS research thesis on improving the efficiency of rotating detonation engines by using aerospike nozzle technologies. A rotating detonation engine (RDE) is a pressure-gain combustion system that utilizes circumferentially-travelling shock waves to detonate the propellants. RDEs are estimated to have specific impulses 10-15% greater than conventional rocket engines.

A propulsion technology with a performance increase of this magnitude is directly applicable to launch vehicles, fulfilling the goal of TABS 1. 2. 4.

Typical RDE engine designs have an annular thrust chamber to facilitate the formation of rotating shock waves. The exhaust flow leaves the combustor through this annulus leading to the formation of a base pressure region behind the center-body of the RDE, which necessitates a nozzle surface within the base region to properly expand the flow and communicate pressure forces to the vehicle.

The annular geometry of the RDE naturally lends itself to utilizing an axisymmetric aerospike nozzle to capture the potential kinetic energy lost in the base region. An aerospike would not only increase the efficiency of the engine, but also allow the engine to have improved performance over a range of pressure altitudes, an additional benefit for a launch vehicle application.

This research project focuses on increasing the efficiency of RDEs by using an aerospike nozzle to further convert the improved energy release from the detonation combustion event into useful kinetic energy. This study would investigate the effect of an aerospike nozzle on the formation of detonation waves in the RDE.

In addition, multiple experiments involving canting the aerospike with respect to the engine centerline will be conducted to determine the ability to thrust-vector the engine. This research is necessary to developing practical rotating detonation engines capable for launch vehicle applications. This research is necessary to developing practical rotating detonation engines capable for launch vehicle applications.

Responsible Mission Directorate Space Technology Mission Directorate (STMD) Responsible Program Space Technology Research Grants (STRG) Lead Organization Purdue University-Main Campus ### **Propulsion Systems** ### 01. 3. 4 Airbreathing Pressure Gain Combustion Applied Research Development Demo & Test Progress in conventional rocket engine technologies, based on constant pressure combustion, has plateaued in the past few decades.

Rotating detonation engines (RDEs) are of particular interest to the rocket propulsion community as pressure gain combustion may provide improvements to specific impulse relevant to booster applications. Despite recent significant investment in RDE technologies, little research has been conducted to date into the effect of nozzle design on rocket application RDEs.

Proper nozzle design is critical to capturing the thrust potential of the transient pressure ratios produced by the thrust chamber. A computational fluid dynamics study was conducted based on hotfire conditions tested in the Purdue V1. 3 RDE campaign.

Three geometries were investigated: nozzleless/blunt body, internal-external expansion (IE-) aerospike, and flared aerospike. The computational study found the RDE's dynamic exhaust plume enhances the ejection physics beyond that of a typical high pressure device. For the nozzleless geometry, the base pressure was drawn down below constant pressure estimates, increasing the base drag on the engine.

For the aerospike geometries, the occurrence of flow separation on the plug was delayed, which has ramifications on nozzle design for operation at a range of pressure altitudes. The flared aerospike design, which has the ability to achieve much higher area ratios, was shown to have potential performance benefits over the limited IE-aerospike geometry. A new test campaign with the Purdue RDE V1.

4 was designed with instrumentation to capture static pressures on the nozzleless and aerospike surfaces. These results were used to validate the results from the computational study. The computational and experimental studies were used to identify new flow physics associated with a rocket RDE important to future nozzle design work.

These results are critical to understanding how to design nozzles for RDEs for both booster and space applications. Recommend changes and additions to this project record.