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 occurrance 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. Future computational work is necessary to explore the effect of different parameters on the nozzle performance. More testing, including with an altitude simulation chamber, would help quantify the possible benefit of new aerospike nozzle designs, including the flared aerospike geometry.

<p dir="ltr">Improving launch vehicle and satellite propulsion system performance directly correlates to the delivery of more mass (or quantity) on orbit from launch vehicles, longer duration satellite missions, and longer ranges for missiles/interceptors. Alternative propulsion devices such as rotating detonation engines (RDEs) offer the potential for significant performance gains but their operability has only been demonstrated on “battle hardened” laboratory devices for rocket applications. The objective of this research was to demonstrate cooling and structural approaches that mature rotating detonation rocket engines (RDREs) to flight like maturation levels.</p><p dir="ltr">Multiple 1.6”/4.1 cm diameter RDE combustors were designed, fabricated, and tested. The RDE tested the most accumulated 309 seconds of hot fire testing and 118 starts/shutdowns. Long duration testing was completed to characterize heat flux and high cycle fatigue (HCF) loading. Large quantities of short duration tests were completed to evaluate thermal cycling impacts to the combustor structure and evaluate low cycle fatigue (LCF) loading. The hardware experienced 118 LCF loadings on the combustor cooling passages, equivalent to the amount of thermal cycle starts and shutdowns. An endurance test was completed at 60 seconds in duration, demonstrating operation well beyond thermal steady state. Additionally, ~3.7 million HCF loadings were placed on the combustor cooling passages, equivalent to the approximate amount of detonation wave passes present for all of the WC 2.0 testing.</p><p dir="ltr">Predicted operating pressures ranged from 5 to 15 atm. The highest-pressure conditions resulted in hot gas wall temperatures exceeding 1000°F on the outerbody of the combustor and injector face temperatures peaking at 350°F. Water calorimetry was used to compute heat fluxes, which were then compared to traditional rocket engine throat level heat fluxes calculated using Bartz equations under average operating conditions. The outerbody heat fluxes reached up to 3.7 kW/cm², while injector face heat fluxes reached a maximum of 1.6 kW/cm². When compared to Bartz throat level values, the outer-body heat fluxes varied from 0.9 to 1.6 times the throat level values, and injector heat fluxes ranged from 0.3 to 0.5 times the throat level values.</p><p dir="ltr">A combined thermal and pressure loading fatigue assessment was completed that took into consideration mean stresses and cumulative damage from the spectrum of loading events. Traditional rocket combustor life is typically limited by the thermal cycles that can be placed on the cooling channel hot wall. The fatigue analysis results highlight the reduction in available low cycle fatigue life as RDE's experience larger thermal loads when compared to traditional rocket combustors. Low cycle fatigue life will become especially challenging in higher chamber pressure combustors where thermal environments are more extreme, and the ability keep hot wall temperatures within acceptable levels is more challenging.</p><p dir="ltr">The study also highlights that the passing detonation wave provides a high cycle fatigue (HCF) failure mechanism that is not present in traditional rocket combustors. This failure mechanism is the result of the pressure pulse provided by the passing detonation wave causing a variable load on the hot wall. This variable load is applied at frequencies commonly in the 10's of kHz, resulting in large quantities of loading cycles when operated at rocket like durations (>60 sec). This HCF failure mechanism is most impactful at larger chamber pressures where the detonation pressure ratio causes peak pressures to be elevated, resulting in larger cyclic stresses and strains in the hot wall. The results indicate that high chamber pressure combustors may experience HCF life exceedances within seconds of operation.</p>