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The issues of Chemical Rocket Engine Control are broadly covered. The basic feedback information and control variables used in expendable and reusable rocket engines, such as Space Shuttle Main Engine, are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control Systems for rocket engines (and vehicles) is presented.
Given here are predictions of fragment velocities and azimuths resulting from the Space Transportation System Solid Rocket Motor range destruct, or random failure occurring at any time during the 120 seconds of Solid Rocket Motor burn. Results obtained using the analytical methods described showed good agreement between predictions and observations for two specific events. It was shown that these methods have good potential for use in predicting the fragmentation process of a number of generically similar casing systems. It was concluded that coupled Eulerian-Lagrangian calculational methods of the type described here provide a powerful tool for predicting Solid Rocket Motor response.
The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.
Design features and concepts that have primary influence on the stable operating flow range of propellant-feed centrifugal turbopumps in a rocket engine are discussed. One of the throttling limitations of a pump-fed rocket engine is the stable operating range of the pump. Several varieties of pump hydraulic instabilities are mentioned. Some pump design criteria are summarized and a qualitative correlation of key parameters to pump stall and surge are referenced. Some of the design criteria were taken from the literature on high pressure ratio centrifugal compressors. Therefore, these have yet to be validated for extending the stable operating flow range of high-head pumps. Casing treatment devices, dynamic fluid-damping plenums, backflow-stabilizing vanes and flow-reinjection techniques are summarized. A planned program was undertaken at LeRC to validate these concepts. Technologies developed by this program will be available for the design of turbopumps for advanced space rocket engines for use by NASA in future space missions where throttling is essential.
An analytical study was conducted to assess and estimate the level of effectiveness of an ejector/rocket or rocket engine nozzle after-burning concept for enhancement of a conventional rocket engine. Performance enhancement and thrust augmentation of an ejector/rocket system were evaluated for a National (or Advanced) Launch System (LNS or ALS) type engine, namely the Space Transportation Main Engine (STME), and its effects on the overall vehicle/propulsion system as payload weight, gross lift-off weight, and propellant weight. The focus was on using a fixed geometry ejector and utilizing the otherwise wasted exhaust excess fuel of rocket engines and burning it with injested atmospheric air to produce additional thrust. Limited analyses were also performed to determine the effect of burning additional injected fuel with the secondary air on thrust augmentation.
Solid rocket propellant and rocket motor components were aged in a vented container on the interior of the LDEF. This paper will present the results of aging the Improved Performance Space Motor-II/Payload Assist Module-Delta II (IPSM-II/PAM-DII) space motor components. Ballistic and mechanical properties of the space aged main propellant, igniter propellant, and ignition system were compared with similar data from preflight and ground aged samples. Mechanical properties of the composite materials and bonded joints used in the motor case, insulation, liner, nozzle, exit cone, and skirt were similarly evaluated. The space aging results will be compared to data collected in a ground based vacuum aging program on similar components. The operation of the vacuum actuated venting valve and pressure actuated resealing of the container will also be addressed. The materials tested showed no significant changes due to space aging. These results indicate that properly designed solid rocket motors can be expected to perform reliably after extended periods of exposure to a space environment.
A fiberoptic Raman diagnostic system for H2/O2 rocket flows is currently under development. This system is designed for measurement of temperature and major species concentration in the combustion chamber and part of the nozzle of a 100 Newton thrust rocket currently undergoing testing. This paper describes a measurement system based on the spontaneous Raman scattering phenomenon. An analysis of the principles behind the technique is given. Software is developed to measure temperature and major species concentrations by comparing theoretical Raman scattering spectra with experimentally obtained spectra. Equipment selection and experimental approach are summarized. This experimental program is part of a program, which is in progress, to evaluate Navier-Stokes based analyses for this class of rocket.
An Intelligent Control System for reusable rocket engines is under development at NASA Lewis Research Center. The primary objective is to extend the useful life of a reusable rocket propulsion system while minimizing between flight maintenance and maximizing engine life and performance through improved control and monitoring algorithms and additional sensing and actuation. This paper describes current progress towards proof-of-concept of an Intelligent Control System for the Space Shuttle Main Engine. A subset of identifiable and accommodatable engine failure modes is selected for preliminary demonstration. Failure models are developed retaining only first order effects and included in a simplified nonlinear simulation of the rocket engine for analysis under closed loop control. The engine level coordinator acts as an interface between the diagnostic and control systems, and translates thrust and mixture ratio commands dictated by mission requirements, and engine status (health) into engine operational strategies carried out by a multivariable control. Control reconfiguration achieves fault tolerance if the nominal (healthy engine) control cannot. Each of the aforementioned functionalities is discussed in the context of an example to illustrate the operation of the system in the context of a representative failure. A graphical user interface allows the researcher to monitor the Intelligent Control System and engine performance under various failure modes selected for demonstration.
The KIVA-2 code, originally developed to solve computational fluid dynamics problems in internal combustion engines, has been developed to solve rocket propulsion type flows. The objective of the work was to develop a code such that both liquid and solid particle motion could be simulated for arbitrary geometry and high speed as well as low speed reacting flows. Modification to the original code include: incorporating independently specific supersonic and subsonic inflows and outflows; symmetric as well as periodic boundary conditions; and the capability to use generalized single or multi-specie thermodynamic data and transport coefficients allowing the user to specify arbitrary wall temperature/heat flux distributions. This code has been shown to successfully solve rocket propulsion flows as well as flows with entrained particles for several different rocket nozzles.
A summary of the Aries experiments and results is presented with specific attention given to the tomographic analysis developed from ground-based photometric data. The instrumentation includes particle detectors, rocket photometers, plasma probes, and 1-deg meridian scanning photometers, and an all-sky TV system. The geophysical conditions during the observations are described including the auroral features and wavelengths, rocket trajectories, and the effectiveness of particle measurements. The tomographic inversion is explained with examples of rocket-photometer height data and contrasted with ground-based photometric data. The modeling techniques used to derive height profiles are described and indirectly compared to measured auroral intensities and intensity ratios. An enhancement of atomic oxygen is inferred from discrepancies with previous studies which confirm the difference between observed intensity ratios and modeled results.
The Liquid Rocket Booster study was initiated by NASA to define an alternative to the Solid Rocket Boosters used on the STS. These studies have involved MSFC, JSC and KSC and their contractors. The prime study contractors, Martin Marietta Corporation and General Dynamics Space Systems, have identified Liquid Booster configurations which would replace the SRB's in the Shuttle stack. The Liquid Rocket Booster increases Shuttle performance to 70 K LBS, provides improved reliability, hold down and verification prior to vehicle release, engine out and improved abort capability, and is phased into the STS launch operations without adversely affecting flight rate.
Hybrid rockets employing liquid oxidizer and solid fuel grain answers to cost, safety, reliability, and environmental impact concerns that have become as prominent as performance in recent years. The oxidizer-free grain has limited sensitivity to grain anomalies, such as bond-line separations, which can cause catastrophic failures in solid rocket motors. An account is presently given of the development effort associated with the AMROC commercial hybrid booster and component testing efforts at NASA-Marshall. These hybrid rockets can be fired, terminated, inspected, evaluated, and restarted for additional testing.
A NASA workshop on rocket-based combined-cycle propulsion technologies is described emphasizing the development of a starting point for earth-to-orbit (ETO) rocket technologies. The tutorial is designed with attention given to the combined development of aeronautical airbreathing propulsion and space rocket propulsion. The format, agenda, and group deliberations for the tutorial are described, and group deliberations include: (1) mission and space transportation infrastructure; (2) vehicle-integrated propulsion systems; (3) development operations, facilities, and human resource needs; and (4) spaceflight fleet applications and operations. Although incomplete the workshop elevates the subject of combined-cycle hypersonic propulsion and develops a common set of priniciples regarding the development of these technologies.
Observations of the UV fluorescent emissions of the NO (1, 0) and (0, 1) gamma bands in the lower-thermospheric dayglow, made with a sounding rocket launched on March 7, 1989 from Poker Flat, Alaska, were analyzed. The resonant (1, 0) gamma band was found to be attenuated below an altitude of about 120 km. A self-absorption model based on Holstein transmission functions was developed for the resonant (1, 0) gamma band under varying conditions of slant column density and temperature and was applied for the conditions of the rocket flight. The results of the model agreed with the measured attenuation of the band, indicating the necessity of including self-absorption theory in the analysis of satellite and rocket limb data of NO.
A technology program aimed at improving the performance of low thrust chemical rockets for spacecraft onboard applications is reviewed. Navier-Stokes analyses of low Reynolds number rocket flows have been compared with local flow property measurements obtained using Rayleigh and Raman diagnostics in a 100 N gaseous hydrogen/gaseous oxygen rocket. It is indicated that computational domain should include the near injector flow and that the shear layer combustion model needs improvement. The system analyses and technical efforts intended to develop a technology base for higher performance propellants are presented. A LOX/hydrazine engine is demonstrated to have a 95 percent theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1.
The goal of the Rocket Based Combined Cycle (RBCC) Propulsion Technology Workshop, was to impart technology information to the propulsion community with respect to hypersonic combined cycle propulsion capabilities. The major recommendation resulting from this technology workshop was as follows: conduct a systems-level applications study to define the desired propulsion system and vehicle technology requirements for LEO launch vehicles. All SSTO and TSTO options using the various propulsion systems (airbreathing combined cycle, rocket-based combined cycle, and all rocket) must be considered. Such a study should be accomplished as soon as possible. It must be conducted with a consistent set of ground rules and assumptions. Additionally, the study should be conducted before any major expenditures on a RBCC technology development program occur.
The major chemical effluents of either solid- or liquid-fueled rockets that can potentially perturb stratospheric ozone include chlorine compounds (HCl), nitrogen compounds (NO(x)), and hydrogen compounds (H2 and H2O). Radicals (Cl, ClO, H, OH, HO2, NO, and NO2) formed directly or indirectly from rocket exhaust can cause the catalytic destruction of ozone. Other exhaust compounds that could presumably lead to ozone destruction either by direct reaction with ozone or by providing a surface for heterogeneous processes include the particulates Al2O3, ice, and soot. These topics are discussed in terms of the possible effects of rocket exhausts on stratospheric ozone.
The Small-scale Solid Rocket Combustion Simulator (SSRCS) program was established at the Marshall Space Flight Center (MSFC), and used a government/industry team consisting of Hercules Aerospace Corporation, Aerotherm Corporation, United Technology Chemical Systems Division, Thiokol Corporation and MSFC personnel to study the feasibility of simulating the combustion species, temperatures and flow fields of a conventional solid rocket motor (SRM) with a versatile simulator system. The SSRCS design is based on hybrid rocket motor principles. The simulator uses a solid fuel and a gaseous oxidizer. Verification of the feasibility of a SSRCS system as a test bed was completed using flow field and system analyses, as well as empirical test data. A total of 27 hot firings of a subscale SSRCS motor were conducted at MSFC. Testing of the Small-scale SSRCS program was completed in October 1992. This paper, a compilation of reports from the above team members and additional analysis of the instrumentation results, will discuss the final results of the analyses and test programs.