Reusable Launch Vehicle Technology Demonstrator Research Articles

Validation of Paras3d Rans Pressure Data with Flight Results Over a Wing-body Rlv Configuration at Supersonic High Angle of Attack Conditions

Winged Reusable Launch Vehicle - Technology Demonstrator (RLV-TD) was successfully flight tested by Indian Space Research Organisation (ISRO). One of the main objectives of this Hypersonic Experiment (HEX) mission is to validate the aerodynamic design and ground based aerodynamic data generation procedure. In this paper, comparison of flight data and CFD estimated pressure coefficient (Cp) over fuselage, wing, vertical tail and rudder are presented at supersonic Mach number and high angle of attack conditions. PARAS3D RANS solver is able to capture the trends as measured in flight. In general, PARAS3D CFD over predicted the compression on windward side and under predicted leeward expansion as compared to flight measurement. In general, CFD predicts Cp close to flight measured value, except over a port in the vertical tail.

ISRO Releases the Special Issue on ‘Reusable Launch Vehicle-Technology Demonstrator’

The Indian Space Research Organisation (ISRO) published the special section on Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) in the 10 January 2018 issue of Current Science.

Development of Large-Sized Titanium Alloy Ti6Al4V and Nickel-Based Superalloy Inconel-718 Forgings for Reusable Launch Vehicle-Technology Demonstrator Flight

Ti6Al4V owing to its high specific strength and Inconel-718 due to its high strength and oxidation resistance at elevated temperatures, have been used in Reusable Launch Vehicle-Technology Demonstrator (RLV-TD). Achieving required properties in large cross-section aerospace quality forgings of these alloys is challenging and has not been reported earlier. Hence, thermo-mechanical processing cycles have been devised to realize forgings of ~1400 mm length in these alloys and successfully used in RLV-TD. Forgings of these alloys having desired microstructures and ultrasonic quality of Class-A1 as per AMS 2630 B standard have been realized. This article discusses the processing challenges and solutions thereof.

Aerothermal Design, Qualification and Flight Performance

Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) experiences severe thermal environment during its ascent as well as re-entry into the atmospheric regime. Structures should be designed to withstand this thermal load. Thermal environments were estimated for RLV-TD and depending on the peak heat flux and heat load, hot structure design for nose cap, wing, vertical tail and control surfaces was developed. Thermal protection system (TPS) using silica tile and flexible insulation was designed to protect the windward and leeward regions respectively. It is essential to verify and establish the design and also physically corroborate the actual thermal performance. Besides the design, the hot structures and TPS functionality as a system has to be qualified and thereby yield full confidence on the total performance during flight. To qualify the hot structures and TPS, various qualification tests were undertaken. The demonstration of their fly-ability and qualification under the combined effect of structural and thermal load was carried out successfully for all structures. This article provides details of aerothermal design of RLVTD and various qualification tests carried out. Comparison of the estimated structure temperatures with measured temperatures in flight shows the robustness of the design methodologies adopted.

Integrated Electrohydraulic Control Actuation System with Centralized Power Plant for the Reusable Launch Vehicle Technology Demonstrator

An electrohydraulic control actuation system for actuating eight aerodynamic control surfaces of the two-stage technology demonstrator of the Indian reusable launch vehicle (RLV-TD) was developed, validated by extensive testing in various test beds, simulation runs and flight proven in the RLV-TD HEX-01 mission flight on 23 May 2016. A centralized hydraulic power generating unit (HPU) was used for powering the eight actuators located in two stages. The network of plumbings and control components of this actuation system were required to be laid out over the entire length and breadth of the vehicle measuring 17 m in length, to distribute the hydraulic power to all the actuators. The hydraulic actuation system had to work for the longest ever duration of 833 s for an Indian launch vehicle. Many challenges arose due to a single HPU for two stages, long complex network of plumbings, longest ever operating duration, thermal issues, second-stage structure resembling an aircraft, limited occupancy for operations at the launch pad, limited power source, etc. Necessary provisions were incorporated in the design to successfully overcome them. This article describes the development of the control actuation system.

An Overview of Reusable Launch Vehicle Technology Demonstrator

After the successful operationalization of Polar Satellite Launch Vehicle and Geo Synchronous Launch Vehicle, the Indian Space Research Organisation is in the process of developing Reusable Launch Vehicle technologies to achieve low-cost access to space. Towards this programme, a winged body configuration was conceived, which can fly at subsonic, supersonic and hypersonic Mach number regime, re-enter into the earth’s atmosphere and simulate the landing manoeuvre. The aerodynamic design, analysis and wind-tunnel testing, aerothermal and structural design, analysis and testing were carried out. Suitable solid motor with slow burn rate propellant was developed. Mission design, guidance and control schemes were implemented. In order to meet the above objectives, certain technologies and infrastructure were developed. The entire subsystems were integrated and a large number of flight measurements were made in the maiden successful flight of Reusable Launch Vehicle Technology Demonstrator in May 2016. The flight measurements and flight performance indicated that the design philosophy, testing schemes and approaches followed are in order, thus providing confidence to proceed to the next logical step in the development of Reusable Launch Vehicle Technologies.

Integrated Navigation, Guidance and Control System and Validation

The navigation, guidance and control (NGC) system of Reusable Launch Vehicle – Technology Demonstrator (RLV-TD) is the most complex system ever flown in ISRO’s launches. It includes a navigation system employing traditional inertial navigation system along with global positioning system aiding and radar altimeter, ceramic servo accelerometer package, angle of attack-based control, different control systems, including secondary injection thrust vector control, reaction control system and aero control surfaces. Closed loop simulations of integrated NGC system of RLV-TD are of utmost importance to ensure mission success by assessment of system performance in nominal and off-nominal flight conditions. Validation is achieved through progressive evaluations in different levels of integrated simulation, namely autonomous simulations, software in loop simulation, on-board processor in loop simulation, hardware in loop simulation and actuator in loop simulation using Iron Bird test facility. This article discusses the overall NGC system, challenges faced during the realization of the simulation test bed, including system design to validate the new on-board elements, commissioning of Iron Bird facility and validation through various phases of simulation. The outcome of these simulations played a crucial role in bringing about improvements and helping in the evolution of the on-board system to the final configuration, ensuring the success of the mission.

Ceramic Coating on Flexible External Insulation Blankets for Reusable Missions

Flexible thermal protection material (flexible external insulation; FEI) is used in the Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) leeward region where multiple curvatures are present. FEI configuration has silica cloth layers on either side with Cerablanket felt sandwiched and stitched together using quartz thread. FEI should ensure an airframe skin temperature in the leeward region within the temperature constraint limits (100°C) of aluminum alloy during the RLV-TD ascent (4.8 W/cm 2 ) and descent (3.7 W/cm 2 ) phases. A ceramic room-temperature coating with low surface solar absorptivity (α) and high thermal emissivity (e) to generate the required thermal gradient at the backwall was developed. The salient features of the ceramic coating are room-temperature curability, ease of application, minimum weight, durability, erosion resistance, stable optical characteristics, easily repairable, non-contaminating, high purity and high temperature compatibility with silica fabric. The properties of the ceramic coating are weight/area: 0.025–0.030 g/cm 2 , solar absorptivity the UV–VIS region: 0.12–0.18, IR emittance: >0.86, backwall temperature: <120°C (from kinetic heating simulation test).

Aerodynamic Design, Characterization and Flight Performance of RLV-TD

Complete aerodynamic design and characterization of ascent as well as descent phases of Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) have been obtained through thousands of wind tunnel tests, flow simulations using various computational fluid dynamic codes and tailor-made engineering codes. The evolution of the ascent and descent designs, aerodynamic characterization and flight performance is presented in this article. The flight data comparison of aerodynamic coefficients, control surface loads, pressure distribution and external acoustic levels with preflight data is good and therefore, providing confidence in using the ground-based data generation techniques along with defined dispersion bands. Some of the minor deviations in the performance observed in flight were resolved/understood in the post-flight analysis, whereas few effects observed in flight need further understanding.

Quality Assurance Challenges in Testing and Evaluation of Reusable Launch Vehicle Systems

Quality assurance of the various reusable launch vehicle subsystems and the certification for flight involve many major and unique challenges compared to the conventional launch vehicle. This article highlights the major challenges faced by quality assurance teams during various phases of realization of the Reusable Launch Vehicle-Technology Demonstrator.

Special Theme: India’s Reusable Launch Vehicle Technology Demonstrator: The Future of Space Transportation System

Special Theme: India’s Reusable Launch Vehicle Technology Demonstrator: The Future of Space Transportation System

Design, Analysis and Qualification of Elevon for Reusable Launch Vehicle

Reusable launch vehicle technology demonstrator is configured as a winged body vehicle, designed to fly in hypersonic, supersonic and subsonic regimes. The vehicle will be boosted to hypersonic speeds after which the winged body separates and descends using aerodynamic control. The aerodynamic control is achieved using the control surfaces mainly the rudder and the elevon. Elevons are deflected for pitch and roll control of the vehicle at various flight conditions. Elevons are subjected to aerodynamic, thermal and inertial loads during the flight. This paper gives details about the configuration, design, qualification and flight validation of elevon for Reusable Launch Vehicle.

RLV-TD Flight Measured Aeroacoustic Levels and its Comparison with Predictions

The Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) is a wing body configuration successfully flight tested. One of the important flight measurements is the acoustic levels. There were five external microphones, mounted on the fuselage-forebody, wing, vertical tail, inter-stage (ITS) and core base shroud to measure the acoustic levels from lift-off to splash down. In the ascent phase, core base shroud recorded the overall maximum at both lift-off and transonic conditions. In-flight noise levels measured on the wing is second highest, followed by fuselage and vertical tail. Predictions for flight trajectory compare well at all locations except for vertical tail (4.5 dB). In the descent phase, maximum measured OASPL occurs at transonic condition for the wing, followed by vertical tail and fuselage. Predictions for flight trajectory compare well at all locations except for wing (− 6.0 dB). Spectrum comparison is good in the ascent phase compared to descent phase. Roll Reaction control system (RCS) thruster firing signature is seen in the acoustic measurements on the wing and vertical tail during lift-off.

Design of Servo Scheme and Drive Electronics for the Integrated Electrohydraulic Actuation System of RLV-TD

Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) is a system which has the ability to carry a payload from the earth’s surface to the outer space more than once. The control actuation forms the major component of the control system and it actuates the control surfaces of the RLV-TD based on the control commands. Eight electro hydraulic actuators were used in RLV-TD for vectoring the control surfaces about their axes. A centralised Hydraulic Power Generating Unit (HPU) was used for powering the eight actuators located in two stages. The actuation system had to work for the longest ever duration of about 850 s for an Indian launch vehicle. High bandwidth requirement from autopilot was met by the servo design using the nonlinear mathematical model. Single Control Electronics which drive four electrohydraulic actuators was developed for each stage. High power electronics with soft start scheme was realized for driving the BLDC motor which is the prime mover for hydraulic pump. Many challenges arose due to single HPU for two stages, uncertainty of aero load, higher bandwidth requirements etc. and provisions were incorporated in the design to successfully overcome them. This paper describes the servo design and control electronics architecture of control actuation system.

Inertial Navigation System for India’s Reusable Launch Vehicle-Technology Demonstrator (RLV-TD HEX) Mission

This work presents the configuration of Inertial Navigation System (INS) used in India’s Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) Program. In view of the specific features and requirements of the RLV-TD, specific improvements and modifications were required in the INS. A new system was designed, realised and qualified meeting the mission requirements of RLV-TD, at the same time taking advantage of the flight heritage attained in INS through various Launch vehicle Missions of the country. The new system has additional redundancy in acceleration channel, in-built inclinometer based bias update scheme for acceleration channels and sign conventions as employed in an aircraft. Data acquisition in micro cycle periodicity (10 ms) was incorporated which was required to provide rate and attitude information at higher sampling rate for ascent phase control. Provision was incorporated for acquisition of rate and acceleration data with high resolution for aerodynamic characterisation and parameter estimation. GPS aided navigation scheme was incorporated to meet the stringent accuracy requirements of the mission. Navigation system configuration for RLV-TD, specific features incorporated to meet the mission requirements, various tests carried out and performance during RLV-TD flight are highlighted.

Quality Assurance and T&amp;E of Inertial Systems for RLV Mission

This work describes the quality assurance and Test and Evaluation (T&E) activities carried out for the inertial systems flown successfully in India’s first reusable launch vehicle technology demonstrator hypersonic experiment mission. As part of reliability analysis, failure mode effect and criticality analysis and derating analysis were carried out in the initial design phase, findings presented to design review forums and the recommendations were implemented. T&E plan was meticulously worked out and presented to respective forums for review and implementation. Test data analysis, health parameter plotting and test report generation was automated and these automations significantly reduced the time required for these activities and helped to avoid manual errors. Further, T&E cycle is optimized without compromising on quality aspects. These specific measures helped to achieve zero defect delivery of inertial systems for RLV application.

Mission Management Computer and Sequencing Hardware for RLV-TD HEX-01 Mission

Reusable Launch Vehicle-Technology Demonstrator Hypersonic Experiment (RLV-TD HEX-01) mission posed some unique challenges in the design and development of avionics hardware. This work presents the details of mission critical avionics hardware mainly Mission Management Computer (MMC) and sequencing hardware. The Navigation, Guidance and Control (NGC) chain for RLV-TD is dual redundant with cross-strapped Remote Terminals (RTs) interfaced through MIL-STD-1553B bus. MMC is Bus Controller on the 1553 bus, which does the function of GPS aided navigation, guidance, digital autopilot and sequencing for the RLV-TD launch vehicle in different periodicities (10, 20, 500 ms). Digital autopilot execution in MMC with a periodicity of 10 ms (in ascent phase) is introduced for the first time and successfully demonstrated in the flight. MMC is built around Intel i960 processor and has inbuilt fault tolerance features like ECC for memories. Fault Detection and Isolation schemes are implemented to isolate the failed MMC. The sequencing hardware comprises Stage Processing System (SPS) and Command Execution Module (CEM). SPS is ‘RT’ on the 1553 bus which receives the sequencing and control related commands from MMCs and posts to downstream modules after proper error handling for final execution. SPS is designed as a high reliability system by incorporating various fault tolerance and fault detection features. CEM is a relay based module for sequence command execution.

Instrumentation and Baseband Telemetry for RLV-TD HEX Mission

In this work, the salient requirements and features of the baseband telemetry system used in Reusable Launch Vehicle—Technology Demonstrator Hypersonic Experiment mission are discussed. The configuration of the overall system, subsystem components and their features are described in brief. The unique requirements of the telemetry system, when compared to that in a conventional launch vehicle, by way of a large number of temperature and strain measurements that enable the assessment of structural integrity and mission performance in re-entry mission, are dealt with, along with the system configuration to cater to these. Subsequently, two new units have been described—Strain Data Acquisition Unit and Multiplexed Data Acquisition Unit that were inducted specifically to cater to strain measurements using strain gauges and temperature measurements using thermocouples respectively. The optimized subsystem configurations for these units are described and their field performance during flight is analyzed. This work further discusses a novel method of data recovery for those measurements affected by the baseline offset shift caused by the presence of a chassis voltage and poor isolation of sensor to chassis.