Title : Radiation measurements in simulated ablation layers

The interaction between an ablating shocklayer and radiative heat transfer to a surface, such as occurs on an ablating hypervelocity re-entry vehicle, is poorly understood. Ablative products act as absorbers and emitters, the net effect of which is strongly dependent on gas composition, enthalpy and wavelength. This paper presents the results of an experimental study investigating radiation in an ablating shocklayer over a 1:13.5 scale Stardust forebody model in an expansion tunnel in air, and nitrogen atmospheres at 9 km/s. The model was coated with a low-pyrolysis temperature hydrocarbon coating, permitting its thermal decomposition and the release of carbon-containing gaseous species into the shocklayer within the 80 μs test time. Shocklayer radiation is visualised using a high speed camera, whilst the emission spectra along the stagnation streamline is measured with ultraviolet and infrared spectrometers. Evidence of coating ablation is presented and shown to produce an increase in radiation in the ultraviolet, in agreement with previous experimental results, and an increase in radiation in the infrared. The influence of oxygen in air is also investigated and found to have a cooling effect on the shocklayer, resulting in a decrease in shocklayer radiation and also shock standoff distance. Objectives The scientific objective of this project was to obtain ablating hypervelocity reentry flows in the laboratory, and to obtain emission spectrometery measurements from the radiating shock layers produced. The data can be used as a benchmark reference for the validation of ablation models and the associated numerical implementations. This will potentially enable spacecraft to be designed with more confidence and safety than before, and to minimize the mass of the thermal protection, without compromising structural integrity. Introduction The design of thermal protection systems (TPS) presents a critical challenge for hypervelocity flight (Gnoffo 1996, Munk 2002). The issue first arose in the 1950‟s when sustained supersonic flight became viable. Early solutions are illustrated by the well known pioneering vehicles the SR71 „Blackbird‟ and the Concorde. These vehicles incorporated „hot structures‟ to balance aerodynamic heat loads, using radiative cooling within the thermal limits of titanium and aluminium respectively at Mach numbers up to 3.2. As flight speed increases, so too does the associated heat input, and more advanced techniques are needed to maintain structural viability. The X15 rocket plane used a combination of inconel and titanium hot structures at speeds up to Mach 7 to survive short durations of hypersonic flight. In parallel with these developments, exploration of space required structures to survive speeds of up to 8 km/sec (Mach 25) for reentry from low earth orbit, and 11.2 km/sec (Mach 35) for return from the moon, and up to 47km/sec for Jupiter entry. Early designs were based on sacrificial ablative heat shields, where aerodynamic heating drives ablative products into the „shock layer‟ of heated gas which surrounds the windward surfaces of hypersonic vehicles. These designs protect the structure of the space craft using the latent heat of the volatiles, through the insulating effect of the vapourised products surrounding the craft, and by radiative cooling from the carbon „char‟ which is left behind after vapourisation. Many products have been successfully used for this purpose, including carbon phenolics (which form the basis of most sacrificial heat shields), cork and modern derivatives such as PICA and CSi coatings (Park 16, Laub 15). Reusable thermal protection systems (TPS) have been developed for speeds of up to 8 km/sec (Mach 25), as exemplified by the US space shuttle. All missions requiring speeds higher than 8 km/s depend on ablative techniques, which form the focus of this application. The mechanisms whereby ablative surfaces provide thermal protection are complex, and cannot be precisely modeled at present (Laux 2009). Current designs contain a large degree of empiricism, and generally use substantial safety factors to account for the unknown parameters. Surface ablation and phase change represent an extreme case of complex coupling between structural and aerothermodynamic boundary conditions, with the interface between structure and fluid flow not being clearly defined. In addition, the ablation proceeds in the environment of a reacting non equilibrium shock layer, which in itself is also hard to characterise. The field lacks accurate experimental data, which makes it difficult to improve our theoretical understanding of the process, and to validate new models. Flight experiments are too expensive and time consuming to use for comprehensive research studies, and program managers are reluctant to include extra scientific instrumentation on reentry capsules as they may compromise the safety of the primary mission. The flight data we do have is very valuable, and remains the subject of study for many decades (Elbert 1993, Cauchon 1967, Corvette 1966, Ried 1972, Capra 2004 and 2007, Na 2008). Ground testing of ablators is difficult. The impulsive facilities which are able to properly simulate the required flow conditions have insufficient run times to fully equilibrate the mass transfer processes involved. For instance, the exposed surface temperatures of an ablator are typically in the range of 2000 to 3000K, and contain a charred layer of carbon, with volatiles diffusing through to the shock layer from a receding and pyrolysing layer underneath. Time scales of the order of 1 to 10 seconds with heating rates in the range of 1 to 100 MWm -2 are required to establish quasi-steady conditions for this system. Arc jet facilities can achieve the required heat transfer rates for sustained periods of time, but do not properly simulate the surrounding hypervelocity flow field, nor its coupling with the ablative process. They are primarily useful for materials testing. Recent developments in The University of Queensland expansion tunnels have demonstrated the ability to create a partial similarity with the flight situation. By exposing a thick layer of epoxy to a hypervelocity flow of 9 km/sec, it has been shown that a flow of ablative products can be established into the shock layer within approximately 20 microseconds. The ablative mixing layer has been probed by spectrometry and high speed photography, and is providing useful data for understanding the process and for evaluation of advanced theoretical models, (D‟Souza 2010). Analysis of the overall mechanisms of ablative protection can be broadly categorized into the following interacting stages and issues:  Materials properties. These are complex and highly temperature and phase dependant, typically involving a composite material consisting of a matrix of fibres, impregnated with a volatile resin. Both the matrix and the volatiles may react chemically with the shock layer.  Heat and mass transfer within the substrate, driven by aerodynamic heating.  The non equilibrium external hypervelocity shock layer flow field.  Diffusion and entrainment of the ablative products into the shock layer.  Radiative processes involving the shock layer, the mixing layer which contains the ablated products, and the exposed surface of the heat shield.  Chemical interaction between shock layer and ablative products The first two processes are best studied using long duration facilities, with conditions set to reproduce the heat transfer rates to be expected in flight. The remaining processes are driven by and coupled to the external flow field, and are the target of this study. By creating a stream of ablative material into the shock layer, the external flow processes can be studied independently of the internal mass and heat transfer mechanisms which are involved in a flight situation. Schematic of processes involved in an ablative TPS shock layer in air. (Laub 2008) Immediately behind the bow shock is a region of non equilibrium chemistry and radiation, which is currently being studied in a related Australian Research Council Discovery project, DP 1094560. At Reynolds numbers of most interest, this region will be physically separated from the ablative mixing layer adjacent to the windward surface of the vehicle. However, it is coupled to it through radiation, and the entrainment of gas which passes from this region into the mixing layer. Uncertainties in this region relate to the post shock chemical kinetics, the associated non equilibrium radiation and the coupling of these processes with the rest of the flow field. The mixing layer itself is characterised by the entrainment of ablated products from the wall into the hypervelocity boundary layer. This may be in the form of vapourised or sublimated volatiles (primarily some form of hydrocarbons), carbon particulates, and a variety of specific additives such as silicon compounds which may be present in the ablator. The mixing layer is also a region of extreme non equilibrium activity, and a participant in radiative processes (either as a transmitter, absorber or source). Hydrocarbons in the mixing layer break down to lighter species thereby absorbing thermal energy, and react with entrained freestream species either endo or exo-thermically. The high carbon content of the ablators leads to the formation of a variety of C species which are typically strong radiators if they get hot enough, such as CN and CO, (Park 2007). Poorly understood processes in the mixing layer include the non equilibrium chemical kinetics and associated radiation, and the influence the mass addition has on the development of the mixing layer and the associated level of turbulence. The rate of mass release is driven by surface heat transfer, which is hard to predict, as it is dominated by radiation and convection through the mixing layer itself. Future spacecraft returning from the Moon, Mars and beyond will re-enter the Earth‟s atmosphere at speeds faster than ever before. The Stardust sample return spacecraft was the fastest successful re-entry of an artificial hypervelocity, entering Earth‟s atmosphere at a speed of 12.8 km/s in January 2006, Jeniskens 2005. Due to its successful re-entry, future hypervelocity sample return spacecraft will be modelled on the Stardust spacecraft. Atmospheric entry at these hypervelocity speeds, typically above 12 km/s, results in extreme heating of the spacecraft. The only thermal protection systems capable of protecting hypervelocity spacecraft are ablative TPS. Whilst the interaction between an ablating TPS and convective heat transfer is understood, little is known of the effect of ablation on radiative heat transfer. As a result, ablative TPSs are typically oversized, sometimes by up to four times, Milos 1997. In order to optimise spacecraft mass to allow maximum mass allocation to the scientific payload whilst providing adequate thermal protection, an improved understanding of the effect of ablation on radiation is desired. No instrumentation was carried onboard the Stardust SRC for the re-entry phase of the mission. The only available flight data was captured from an airborne observatory, and is limited to spectra of the entire shocklayer without spatial resolution, Winter 2007. Although not verified either in flight or ground tests, several computational predictions have also been developed, Johnston 2007,Olyniack 1999, Park 2007. Present work consists of ground testing of radiation from ablating shocklayers in hypersonic flows to enable verification of computational codes. The ground-based test facilities most closely able to replicate the high enthalpy, chemically reacting shocklayers characteristic of super-orbital re-entry flows are expansion tube facilities, Morgan2006. These are impulse facilities with brief, steady test periods of the order of 50 to 500 μs. However, ablation is a chemical and physical phenomenon requiring prior heating, previously considered to be of the order of minutes. Outgassing associated with ablation is known to contaminate the test conditions within an expansion facility, thereby rendering ablation simulation in expansion facilities difficult. Ablation simulation in expansion facilities had therefore been limited to a subset of ablative heat shield phenomena, such as gaseous product effusion from the forebody surface during the brief steady test time, Morgan 1999, and simulation of a pre-heated graphite layer, Hunt 2002. Proof-of-concept experiments have demonstrated that a low-pyrolysis temperature hydrocarbon (epoxy) coating may be used to study ablation in impulse facilities, D‟Souza 2010. Radiating, carbon-containing ablative material, arising from pyrolysis of an epoxy coating, is used to simulate an ablating TPS. This paper presents a summary of the experimental work conducted in expansion facilities at The University of Queensland on the interaction between an ablating shocklayer and radiative heat transfer around a Stardust scale model forebody. Test gases used are air and nitrogen. The epoxy coating employed in this work is „Five Minute Araldite‟ (®Huntsman Corporation). High speed camera and spectrometric results show ablation of the coating, in agreement with previous findings. The results of investigations into the effect of various test gases with wavelength are presented here. ANALYSIS A 1D semi-infinite analysis was performed to demonstrate that vaporisation of an epoxy coating is possible within the brief test times available in expansion facilities. Conditions taken in the analysis of the heat flux to the epoxy coating are given in Table 1. The Zoby stagnation point heat flux approximation for convective heat flux to the (cold) nose of the model (Eq. 1, Zobt 1968) is used in conjunction with the Schultz and Jones temperature change for a constant heat flux during the exposure time (Eq. 2, Schultz and Jones 1973):