A MHz-rate mid-infrared laser absorption sensing strategy has been developed to simultaneously measure the intra-cycle variation of carbon monoxide (CO), carbon dioxide (CO2), 2 ), water (H2O), 2 O), pressure, and temperature in the chamber of rotating detonating rocket engines (RDREs). The beams of two quantum cascade lasers (QCL) and one interband cascade laser (ICL) were multiplexed to target rovibrational transitions in the fundamental bands of CO and H2O 2 O (at 5 mu m) ) and CO2 2 (at 4.2 . 2 mu m ). Extended trapezoidal injection current modulation via bias- tee circuitry provided sufficient spectral range to fully resolve collisionally broadened spectral transitions with a 1-mu s mu s time resolution. The three laser beams are coupled into a single optical fiber for light delivery through a sapphire window and into the annulus of a gaseous methane-oxygen rotating detonation rocket engine (RDRE) at the Air Force Research Lab in Edwards, CA. The light was retro-reflected from the annulus centerbody and spectrally demultiplexed onto two photovoltaic (PV) detectors. Quantitative, three-species concentration measurements, along with temperature and pressure, were inferred from the spectrally-resolved signals at 1 MHz over a range of post-detonation, in-chamber conditions, 1500-3000 K and 1-5 atm at targeted propellant mass flow rates ranging from 0.09-0.36 kg/s. The integrated five-parameter measurement scheme captures approximately 80% of the total combustion gas composition and fully characterizes the thermodynamic state at detonation-relevant time-scales, offering broad applicability to high-speed combustion flows. Novelty and significance statement This work involves the development and application of a novel five-parameter, sub-microsecond combustion diagnostic for highly dynamic gas properties in the chamber of a rotation detonation rocket engine. The laser spectroscopy method is uniquely comprehensive in defining thermodynamic state from simultaneous species (CO, CO2, 2 , and H2O), 2 O), temperature, and pressure measurements attained using a multiplexed mid- infrared laser absorption strategy. The diagnostic captures 80-90% of the species composition, which reduces compositional uncertainty in collisional broadening and thereby provides a highly quantitative MHz-rate optical pressure measurement. The quantitative, time-resolved pressure measurement is critical in the assessment of pressure-gain combustion systems, wherein traditional pressure transducers are largely inadequate, while the species and temperature measurements also serve as important reference data for simulations supporting the community's efforts to advance rotating detonation engine technology.

A MHz-rate mid-infrared laser absorption sensing strategy has been developed to simultaneously measure the intra-cycle variation of carbon monoxide (CO), carbon dioxide (CO2), water (H2O), pressure, and temperature in the chamber of rotating detonating rocket engines (RDREs). The beams of two quantum cascade lasers (QCL) and one interband cascade laser (ICL) were multiplexed to target rovibrational transitions in the fundamental bands of CO and H2O (at 5µm) and CO2 (at 4.2µm). Extended trapezoidal injection current modulation via bias-tee circuitry provided sufficient spectral range to fully resolve collisionally broadened spectral transitions with a 1-µs time resolution. The three laser beams are coupled into a single optical fiber for light delivery through a sapphire window and into the annulus of a gaseous methane-oxygen rotating detonation rocket engine (RDRE) at the Air Force Research Lab in Edwards, CA. The light was retro-reflected from the annulus centerbody and spectrally demultiplexed onto two photovoltaic (PV) detectors. Quantitative, three-species concentration measurements, along with temperature and pressure, were inferred from the spectrally-resolved signals at 1 MHz over a range of post-detonation, in-chamber conditions, 1500–3000 K and 1–5 atm at targeted propellant mass flow rates ranging from 0.09–0.36 kg/s. The integrated five-parameter measurement scheme captures approximately 80% of the total combustion gas composition and fully characterizes the thermodynamic state at detonation-relevant time-scales, offering broad applicability to high-speed combustion flows.Novelty and significance statementThis work involves the development and application of a novel five-parameter, sub-microsecond combustion diagnostic for highly dynamic gas properties in the chamber of a rotation detonation rocket engine. The laser spectroscopy method is uniquely comprehensive in defining thermodynamic state from simultaneous species (CO, CO2, and H2O), temperature, and pressure measurements attained using a multiplexed mid-infrared laser absorption strategy. The diagnostic captures 80-90% of the species composition, which reduces compositional uncertainty in collisional broadening and thereby provides a highly quantitative MHz-rate optical pressure measurement. The quantitative, time-resolved pressure measurement is critical in the assessment of pressure-gain combustion systems, wherein traditional pressure transducers are largely inadequate, while the species and temperature measurements also serve as important reference data for simulations supporting the community’s efforts to advance rotating detonation engine technology.

As scaling of rotating detonation rocket engines becomes increasingly important for transition from laboratory-scale experiments to application-based testing and flight demonstrations, scaling methodologies must be studied and tested. In this study, a 76.2~mm RDRE is scaled up to a 101.6~mm annulus, proportionally scaling the injection area such that the ratio of the annulus area to the injection area is constant. This enables similar injector response for both geometries. Additionally, as symmetric injector response has been shown to be important to RDRE operation, a new injector is designed that provides axial net-momentum balance and symmetric plenum pressures around phi=1.2. These three geometries are compared on the basis of their operability limits, system pressures, detonation wave propagation and global performance. It is shown that the scaling methodology used produces an engine that provides similar performance at equivalent total mass flux, and also demonstrates that the momentum-balanced injector with greater injection area provides the same global performance with significantly reduced plenum pressures and increased detonation wave speed.

View Video Presentation: https://doi.org/10.2514/6.2023-0355.vid Additively manufactured, water-cooled test hardware was developed for calorimetry analysis in a rotating detonation rocket engine architecture. The water-cooled chamber hardware was designed for extended-length hot-fire testing at the Air Force Research Laboratory (Rocket Propulsion Division) in Edwards, CA, and features sensors to recover local heat fluxes at locations of interest from the injector face and global heat transfer to the combustor walls. A design leveraging metal additive manufacturing was utilized to generate cooling channels to "snake" around sensor ports, enabling heavy instrumentation of a monolithic, axially cooled chamber without compromising coolant flow for traditional subtractive manufacturability. Conjugate heat transfer analyses (CHT) were performed to predict the temperatures within the chamber wall of the test hardware at steady state conditions for a given coolant flow rate and to determine the impact of channel snaking on local heat transfer and thermocouple placement. The calorimetry hardware developed is a critical step in demystifying the otherwise convoluted heat transfer mechanisms within rotating detonation rocket engine architectures, and will enable long-duration hot-fire tests hitherto unachievable with traditionally-manufactured rotating detonation rocket engine hardware designs. With the design complete, manufacture of the chamber is anticipated to be completed by the end of Winter 2023, and hot-fire testing and performance characterization of the chamber are planned to be performed at AFRL-Edwards in Spring 2023. Subsequent measurements of temperature and heat flux during hot-fire operation are expected to provide unique insights into the practical application of detonation-based thermodynamic cycles for rocket propulsion research and development.

Rotating detonation rocket engine (RDRE) experiments have so far been limited to injectors with high pressure drop to limit injector-plenum coupling. However, this corresponds to a reduction in the potential performance benefits of RDREs. To investigate this phenomenon, two injection configurations with different total injection areas are compared with respect to overall engine operability, global performance, and steady-state wave propagation. Notably, the increased area corresponds with a three to four times decrease in pressure drop, with negligible effects on operability and performance. The RDRE is found to operate in a detonative mode for both injectors across equivalence ratios phi between 0.6 and 2.5, and total mass flows (m) over dot(tot) between 0.09 and 0.45kg/s. The steady-state wave propagation varies significantly between the two injectors, with low pressure drop injection sustaining fewer waves traveling at higher velocities at the maximum performance conditions, but in all other conditions promoted lower wave speeds and counter-propagating behavior, denoting a breakdown in the steady wave propagation. This breakdown is likely due to increased injector-plenum coupling. These results provide insight into the effects of reduced pressure drop on the operation of an RDRE, and indicate challenges that must be overcome in the design of real systems that include RDREs.

A rotating detonation rocket engine (RDRE) with various convergent nozzles and chamber lengths is investigated. Three hundred hot-fire tests are performed using methane and oxygen ranging from equivalence ratio equaling 0.5–2.5 and total propellant flow up to 0.680 kg/s. For the full-length (76.2 mm) chamber study, three nozzles at contraction ratios ϵc = 1.23, 1.62 and 2.40 are tested. Detonation is exhibited for each geometry at equivalent conditions, with only fuel-rich operability slightly increased for the ϵc = 1.62 and 2.40 nozzles. Despite this, counter-propagation, i.e., opposing wave sets, becomes prevalent with increasing constriction. This is accompanied by higher number of waves, lower wave speed Uwv and higher unsteadiness. Therefore, the most constricted nozzle always has the lowest Uwv. In contrast, engine performance increases with constriction, where thrust and specific impulse linearly increase with ϵc for equivalent conditions, with a 27% maximum increase. Additionally, two half-length (38.1 mm) chambers are studied including a straight chamber and ϵc = 2.40 nozzle; these shortened geometries show equal performance to their longer equivalent. Furthermore, the existence of counter-propagation is minimized. Accompanying high-fidelity simulations and injection recovery analyses describe underlying injection physics driving chamber wave dynamics, suggesting the physical throat/injector interaction influences counter-propagation.

Time-resolved in situ measurements of thermodynamic properties (pressure, temperature and species density) were performed in the annular exhaust of a methane-oxygen rotating detonation rocket engine (RDRE) using high-speed laser absorption spectroscopy. Bias-tee circuitry was coupled with a distributed feedback quantum cascade near 5 μm to spectrally-resolve a cluster of rovibrational transitions in the fundamental vibrational band of CO at MHz measurement rates, from which temperature and density were inferred from the line areas and pressure from line-width. The laser source was fiber-coupled for remote light delivery to the exhaust plane of the RDRE combustor. A single-ended optical interface with retro-reflection off of the annulus center body was utilized for stand-off in situ detection. The test article consisted of a 76.2 mm diameter annulus with a annular width of 0.5 cm and doublet impingement injection. Time-resolved CO absorption data was analyzed to examine the evolution of gas properties during engine test firings of 0.5–1.0 s in duration. Start-up transients were examined and intra-cycle profiles of gas properties were compared over a range of equivalence ratios and mass fluxes comprising approximately 25 test conditions. The effects of variable mixing were also examined through a staggered doublet injector configuration. The novel thermodynamic dataset was compared with thrust measurements, wave speed visualizations and first-order detonation models, revealing non-ideal behavior such as parasitic deflagration and modulating oxidizer-to-fuel ratio related to post-detonation injector recovery.

We present the methodology and implementation of best practices developed at the Air Force Research Laboratory’s Rocket Laboratory for metering both gaseous and liquid propellant flows in a rotating detonation rocket engine (RDRE). RDREs produce high-frequency and large amplitude pressure waves in the combustion chamber which create an upstream dynamic response in the propellant supply. As detonation behavior is highly dependent on equivalence ratio and propellant mixing, it is desirable to limit the impact of these back-pressure events on propellant flow. For gases, a sonic nozzle – which uses an area constriction to create sonic flow– is advised as it simultaneously isolates the propellant flow rate from downstream pressure fluctuations and enables metering of the flow. With proper implementation, uncertainty in mass flow rate measured with a sonic nozzle will be on the order of 2%, primarily driven by the uncertainty in the temperature measurement. For liquids, flow isolation and metering can be achieved using a cavitating venturi. Uncertainty in flow rate measured with a cavitating venturi should be less than 1% and is primarily driven by uncertainty in propellant vapor pressure and the calibrated discharge coefficient. If followed, this guidance to the community will ensure not only low uncertainty in reported results, but commonality of approach and greater synergy between research efforts.

A mid-infrared laser absorption sensing method has been developed to quantify gas properties (temperature, pressure, and species density) at MHz measurement rates, with application to annular rotating detonation rocket flows. Bias-tee circuitry is integrated with distributed feedback quantum cascade and interband cascade lasers in the $$4{-}5~\mu \hbox {m}$$ range enabling diplexed radio frequency (RF) wavelength modulation on the order of several MHz while yielding sufficient scan depth to capture multiple rovibrational transitions in the fundamental vibrational bands of $${\text {CO}}$$ and $${\text {CO}}_{2}$$ . Sub-microsecond spectrally-resolved $${\text {CO}}$$ absorption lineshapes provide for inference of temperature and species from a two-line area ratio and pressure from collision line-width. $${\text {CO}}_{2}$$ column density is inferred from peak-to-valley differential absorption at the bandhead near $$4.19~\mu \hbox {m}$$ . A field demonstration on a methane-oxygen rotating detonation rocket engine was performed utilizing an in situ single-ended retro-reflection optical configuration aligned at the exhaust plane. The target gas properties are temporally-resolved at up to 3 MHz across rotating detonations with up to 20 kHz cycle frequency.

The paper presents experimental evidence of continuous detonation in a rotating detonation rocket engine (RDRE) powered by H2/O2 propellants. High-speed chemiluminescence imaging is used to characterize the detonation wave dynamics by introducing a tracer in the hydrogen fuel flow. The results show continuous five-wave co-rotating detonations at various equivalence ratios and flow rates demonstrating the potential for H2/O2 propellant based RDREs for upper-stage rocket engines.

A unique spectroscopic strategy has been developed for laser absorption sensing of carbon monoxide (CO) and carbon dioxide (CO2) at extreme pressures (P > 50 atm) relevant to modern combustion devices. The strategy exploits the band narrowing effects of line mixing, which acutely impact spectrally dense regions, such as bandheads, where line spacing is small. Line mixing is shown to counter collisional line-broadening effects that reduce differential absorption at elevated pressures and often limit the pressure range of laser absorption methods. In this work, the R-branch bandheads of CO and CO2, which are only observed at high temperatures relevant to combustion, are targeted near 2.3 mu m and 4.2 mu m, respectively. Spectral line-mixing models were developed for each bandhead region to account for the collision-induced population transfer rates between rotational energy states over a wide range of elevated temperatures and pressures. Modified exponential-gap models using the relaxation matrix formalism were shown to capture the thermodynamic dependence of the population transfer rates and enabled scaling. Differential absorption at the bandheads was observed to increase by up to a factor of ten at high gas densities, due to line-mixing effects, enabling detection with relatively narrow-band tunable semi-conductor lasers. With refined spectroscopic models, laser absorption measurements of temperature, CO, and CO2 were demonstrated over a range of high pressures (up to 104 atm) in a sub-scale rocket combustor operated with kerosene and supercritical methane. (C) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

A novel laser absorption sensing strategy has been developed to evaluate combustion progress through quantitative measurements of carbon dioxide (CO2) in high-pressure (> 50atm), high-temperature (> 3000 K) hydrocarbon-fueled rocket combustion flows. The sensor enables a broad range of operability by probing rovibrational transitions in the bandhead of CO2 near 4.2μm, accessed with an interband cascade laser. Under extreme rocket conditions, this targeted bandhead region experiences line-mixing effects that favorably distort the molecular spectra. A preliminary spectroscopic model of line-mixing effects has been developed utilizing a high-enthalpy shock tube to achieve scalability of spectral simulations over a range of high temperatures and high pressures. The model is employed for quantitative interpretation of measured absorption signals. The mid-infrared light source was fiber-coupled for remote light delivery at propulsion test facilities. A wavelength modulation spectroscopy technique utilizing normalized-second harmonic detection was implemented for acquiring differential absorption signals in a harsh rocket combustor environment. Using this method, measurements of CO2 concentration have been demonstrated over a range of operating conditions up to 83 bar in a single-element-injector RP-2/GOx rocket combustor at the Air Force Research Laboratory in Edwards, CA.

This paper presents analysis of results from optical diagnostics in a high-pressure combustor burning gaseous oxygen (GOX) and liquid kerosene RP-2 fuel through a jet-swirl coflow injector. The objectives of the experiment were to measure flame stabilization and position under high-pressure conditions. Data were obtained at pressures from 2 to 16.5 MPa and mixture ratios from 2.9 to 20. High-speed cameras captured side-on chemiluminescence and infrared images of the flame. Results show that the flame-spreading angle from the injector ranges from approximately 3 to 6 deg, varying with pressure and propellant mass flow rate. A novel borescope was used to image the flame from upstream of the GOX post, enabling visualization of the flame near its stabilization location. The flame stabilization characteristics change significantly with the fuel flow velocity (which is proportional to pressure). At low pressures and fuel flow velocities, the flame appears to be distributed azimuthally in a nearly axisymmetric manner. At higher velocities, helical spirals of luminosity develop near the GOX post. In addition, the side-on views reveal concentrated streaks of, presumably, fuel entering the combustion chamber. These results suggest that computations must resolve the individual fuel injection orifices, to capture the flame stabilization.

A novel cross-band laser absorption spectroscopy technique has been developed for quantitative measurements of gas temperature and carbon monoxide (CO) in high-pressure, high-temperature rocket combustion flows. The strategy enables a broad range of sensor operability by simultaneously probing rovibrational transitions in both the fundamental and first overtone bands of CO near $$4.98\,\upmu \hbox {m}$$ and $$2.32\,\upmu \hbox {m}$$, respectively, which sustain large differences in temperature dependence despite collisional broadening. Scanned-wavelength modulation spectroscopy methods are integrated for noise rejection in the harsh rocket operating environment. Initial experiments using the cross-band thermometry technique have been conducted on a single-element-injector rocket combustor with RP-2/GOx and $${\hbox {CH}_4}/\hbox {GOx}$$ propellant combinations at pressures up to 75 bar. Measurements of the first overtone bandhead ($$2.32\,\upmu \hbox {m}$$) maintained adequate signal-to-noise at even higher pressures (up to 105 bar), although deviating significantly from spectral simulations. To account for collisional effects at high gas densities, empirical models for line mixing, developed via shock tube studies, were employed to enable quantitative interpretation of measured signals for temperature and CO mole fraction in the rocket combustor.

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/hydrocarbon propellant multi-element integrated test article that included an oxidizer-rich oxygen/hydrocarbon propellant preburner and a staged-combustion main injector. Also as part of this project, the Air Force Research Laboratory fabricated single-element main injectors of the same designs as used in the NASA multi-element injectors, and tested them in a staged-combustion integrated test article that used an oxidizer-rich oxygen/hydrogen propellant preburner. Final results of the multi-element and single-element staged-combustion main injector test programs are described in companion papers at this JANNAF meeting. The design, development, and preliminary test results of these main injectors have also been described in previous JANNAF papers. The main injector element designs were all based on relatively conventional gas-centered swirl coaxial injector element configurations such as used in Russian RD-170 and NK-33 engines, and planned for use in future U.S.-built experimental engine systems such as the Hydrocarbon Boost program demonstration engine. Four different elements were tested in both the multi-element and single-element main injectors, at similar combustion chamber pressures, chamber contraction ratios, and mixture ratios. Variations of the element features included recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. This paper compares the hydraulics, combustion performance, stability, and compatibility characteristics of the single-element and multi-element injectors operated at similar conditions. The single-element hardware is shown to have captured a significant level of the operability of the multi-element hardware.