Ight Experiment Research Articles

Thermal Modeling for the Sustained Hypersonic Flight Experiment

This paper details the development, validation and use of a thermal model created for the UK Ministry of Defence funded Sustained Hypersonic Flight Experiment (SHyFE). This compact ramjet will operate at Mach 4 to 6 leading to very high air ∞ow stagnation temperatures. The thermal management system must be designed to maintain the avionics and fuel system at acceptable temperatures throughout the ∞ight. The heat exchange mechanisms were predicted throughout the acceleration section of the Onal SHyFE ∞ight experiment using a thermal resistance network approach which incorporates convection, radiation and conduction. This allowed the SHyFE thermal management system to be assessed at an early design stage before the vehicle geometry was Onalised and before detailed CAD models were available. The resulting thermal model is highly ∞exible which allows changes in geometry to be easily incorporated and assessed. The thermal model of the SHyFE vehicle during ∞ight was validated using SHyFE Mach 4 combustion ground test surface temperature measurements. The ∞exibility of the model allowed the combustion chamber and nozzle ground experiment thermal systems to be predicted using the same methodology as the ∞ight experiment. The predicted combustor exit gas temperatures show good agreement with the measured ground experiment exit gas temperatures providing conOdence in the thermal modelling technique. The Mach 4 combustion test thermal models were used to aid the design of the ground experiment cooling water system and to assess the reliability of experiment high temperature instrumentation. The SHyFE ∞ight thermal model was used to assess the suitability of a range of ∞ameholder designs by predicting the nozzle heat ∞ow and surface temperatures for each design.

Simulation of Deployment Dynamics of Inflatable Structures

Analytical simulation of the ine ation process of ine atable structures is key to assessing their robust deployment in a space environment. Although the problem seems to be intractable, a simplie ed e nite volume ine ation model is developed, which allows discretized description of the gas e ow, pressure variation, and resulting nonlinear large deformation throughout domains of the ine atable structure. Simulation results are presented along with convergence studies, which provide insight into the deployment of three different packaging/deployment schemes: Z-unfolding, rollout, and extrusion of a cylindrical tube. The proposed ine ation model is not limited to prismatic one-dimensional components and can be applied to ine atable shapes with complex cone gurations. As a validation of the numerical simulation, a laboratory experiment was conducted on the Z-folded cone guration. Generally, good agreement in the overall ine ation dynamics is observed between the analysis and experiment. EVERALspacemissionshaveconsideredtheuseoflightweight ine atable structures for components such as booms, sunshades, solar concentrators, solar sails, and antennas for nearly all aspects of Earth and space explorations. As a prelude to these missions, the ine atable antenna experiment (IAE) was deployed from the Space Shuttle Endeavour in May 1996 to demonstrate the readiness and reliabilityoftheine atabletechnologyforalarge14-mantennastructure in a realistic space environment. 1 One of the urgent technology issues revealed by the brief 80-min IAE e ight experiment was the need to better understand the dynamics of deploying ine atable structuresinspaceandhowthe ine ationprocessisine uenced by the deployment scheme. This includes the initial packaging and subsequent release and ine ation. Depending on the deployment scheme, large ine atable structures can be extremely e exible to the point of instability, especially during the early stages of ine ation. Rigidization can begin only after e nal deployment is achieved. Testing the deploymentof alargeine atableinan Earth environmentin the presence of gravity and air is of limited value for inferring its deploymentbehavior in space. Therefore, analytical models for simulating and predicting the dynamics of the ine ation process are essential tools for understanding ine atables’ deployment behavior in space environment and for guiding and improving future packaging and deployment concepts. To thisend,this paper explores theuse ofa e nite volume ine ation model that allows a tractable simulation of the deployment process of various ine atable cone gurations, from their initial stowed state to full deployment. Of interest here is a description of all states of ine ation as a function of time. Knowledge of all states of ine ation is essential for subsequent assessment of conditions of stability and controllability of the deployment process.

Simulator Platform Motion Effects on Pilot-Induced Oscillation Prediction

Simulator motion platform characteristics were examined to determine if the amount of motion affects pilotinduced oscillation prediction. Five test pilots evaluated how susceptible 18 different setsof pitch dynamics were to pilot-induced oscillations with three different levels of simulation motion platform displacement: large, small, and none. The pitch dynamics were those of a previous in-e ight experiment, some of which elicited oscillations. These in-e ight resultsserved as truth data for the simulation. As such, the in-e ight experiment wasreplicated asmuch as possible. Objective and subjective data were collected and analyzed. With large motion, pilot-induced oscillation and handling qualities ratings matched the e ight data more closely than with small motion or no motion. Also, regardlessof the aircraft dynamics, large motion increased pilot cone dencein assigning handling qualitiesratings, reduced safety pilot trips, and lowered touchdown velocities. Whereas both large and small motion provided a pitch rate cue of high e delity, only large motion presented the pilot with a high e delity vertical acceleration cue.

Classically Rationalized Low-Order, Robust Structural Controllers

An experimental and analytical comparison of classically rationalized and optimal control design techniques is conducted. Classically rationalized compensators blend the loop assignments, complex topological design, and robustness of optimal controllers and the lower order, robustness, and practical insight of classical controllers. The single-input/single-output disturbance rejection problem for controlled structures is divided into four distinct topologies, and optimal controllers aredesignedand interpreted classically.Theobservationsarethen summarized into a set of design rules forthedesign oflow-order robustcontrollers. Classically rationalized and robust compensators are experimentally implemented on the Middeck Active Control Experiment, a Shuttle e ight experiment

On-Orbit Robust Control Experiment of Flexible Spacecraft ETS-VI

An on-orbit attitude and vibration control experiment has been performed as a part of experiments of the ETS-VI, which was launched by the H-II rocket in August 1994. The purpose is to develop precise attitude control technology for future large spacecraft or large space structures. To this end, we have designed three types of controllers in linear quadratic Gaussian and H 1 frameworks to achieve the robust stability against residual modes and modal parameter errors. The experiments were done at the beginning of 1995, and their feasibility has been proved. The  ight experiment results and the modeling and the controller synthesis methods are reported. Based on the on-orbit results, the capabilities of controllers are compared and discussed.

Uncertainty Assessment of Performance Evaluation Methods for Solar Thermal Absorber/Thruster Testing

Proposed ground testing of solar thermal upper stage propulsion systems will verify the technology readiness level and provide a higher degree of cone dence for pursuing a e ight experiment. This study applies experimental uncertainty analysis techniques to investigate various performance evaluation methods that could be used to interpret results of solar thermal absorber/thruster ground testing. A general uncertainty analysis is applied to engine specie c impulse determination methods for an open-ended cavity, heat exchanger type, solar thermal engine. Six performance relationships consisting of both thrust- and nonthrust-based evaluation methods are utilized and two nominal operating cases are assessed. Based on the uncertainty estimates made for individual variables in the data reduction equations, overall specie c impulse uncertainties for the nonthrust-based methods range from about 6 to 10%, depending on the method and testing situation. Thrust-based evaluation methods yield specie c impulse uncertainties on the order of 1.5% if thrust and propellant mass e ow rate can be measured with uncertainties of 1% or less.

Aerodynamic Characteristics Evaluation of Hypersonic Flight Experiment Vehicle Based on Flight Data

A hypersonic e ight experiment was conducted for development of the Japanese unmanned operational orbiting plane. The following aerodynamic characteristic data are presented for a e ight ranging in Mach number from 2 to 14: aerodynamic force coefe cients, longitudinal trim characteristics, stability and control derivatives, elevon hingemoment, and surfacepressuredistribution.Thedata arecomparedwithpree ightpredictionsbased on windtunnel test results and computational e uid dynamiccalculations to evaluate the validity of the prediction methods, including determination of uncertainty in the vehicle design. The prediction methods are proven to be generally valid for design of lifting re-entry vehicles at a high angle of attack, whereas some anomalies are found in axial force, elevon trim angle, and reaction-control-system gas-jet interaction.

Experimental Investigation of Spacecraft In-Flight Disturbances and Dynamic Response

In September 1991, NASA launched the Upper Atmosphere Research Satellite. In addition to its atmospheric sciences mission, analysis of data from the e rst 370 days after launch was used to investigate in-e ight spacecraft disturbances and responses. The investigation included a three-orbit in-e ight experiment to determine how each onboard instrument and subsystem disturbance contributed to the overall spacecraft dynamic response. The investigation quantie ed the spacecraft dynamic response produced by the solar array and high-gain antenna harmonic drive disturbances. The solar array’ s harmonic drive output resonated two solar array modes. Friction in the solar array gear drive provided sufe cient energy dissipation, which prevented the solar panels from resonating catastrophically; however, the solar array vibration amplitude was excessively large. The resulting vibration had a latitude-specie c pattern. Thermal elastic bending of the spacecraft’ s two e exible appendages as the spacecraft crosses the Earth’ s terminator and solar array modal contribution to the spacecraft response were also examined.